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Full text of "Apollo 13 - Press Kit"

OX-. 1 -? -^ 



NEWS 



MA^A 



RELEASE NO: 70-50K 



NATIONAL AERONAUTICS ANO SPACE ADMINISTRATION 
WASHINGTON, D .C . 20546 

FOR RELEASE: Thursday a.m. 

April 2, 1970 



TEIS. 



WO 2-4155 
WO 3-6925 



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COPY 



mi 2 wo 



PROJECT: 




APOLLO 13 




K 

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contents 

GENERAL RELEASE 1-9 

APOLLO 13 COUNTDOWN 10-12 

Lightning Precautions — ■ 12 

May Launch Opportunities 13 

Apollo 13 Flight Profile 14 

LAUNCH, MISSION TRAJECTORY AND MANEUVER DESCRIPTION 15 

Launch 15-16 

Launch Events 17 

Apollo 13 Mission Events 18-23 

Earth Parking Orbit (EPO) 24 

Translunar Injection (TLI) 2 4 

Transposition, Docking, and Ejection (TD&E) 24 

Saturn Third Stage Lunar Impact 24-26 

Translunar Coast 26-27 

Lunar Orbit Insertion (LOI) 27 

Descent Orbit Insertion (DOI) 27 

Lunar Module Separation :-27 

CSM Circular! zat ion 2 7-28 

Power Descent Initiation (PDI), Lunar . Landing 29 

Lunar Surface Exploration 29-37 

Ascent, Lunar Orbit Rendezvous 37-41 

Ascent Stage Deorbit 42 

Transearth Injection (TEI) 42 

Transearth Coast 42-43 

Recovery Operations 44-45 

APOLLO 13 ONBOARD TELEVISION 46 

Apollo 13 TV Schedule 47 

APOLLO 13 SCIENCE 4 8 

Lunar Orbital Photography 48-49 

Charged Particle Lunar Environment Experiment (CPLEE) 49-51 

Lunar Atmosphere Detector (LAD) 51-53 

Lunar Heat Flow Experiment (HFE) 54-58 

Passive Seismic Experiment (PSE) 58-60 

Solar Wind Composition 60-61 



-. /n s 



•62 



APOLLO 13 SCIENCE (Cont'd.) 

Dust Detector 

Field Geology Investigations 63-65 

SNAP-27 °5-67 

PHOTOGRAPHIC EQUIPMENT bbi-by 

LUNAR. DESCRIPTION 70 

Physical Facts 70 

Landing Site 71-72 

APOLLO 13 FLAGS, LUNAR MODULE PLAQUE 7 3 

SATURN V LAUNCH VEHICLE 74 

First Stage 74 

Second Stage 74 

Third Stage ^ 75 

Instrument Unit 76 

Propulsion — 76-77 

COMMAND AND SERVICE MODULE STRUCTURE, SYSTEMS 78 

Launch Escape System (LES) 78 

Command Module (CM) Structure 78 

Service Module (SM) Structure 78-80 

Spacecraft-LM Adapter (SLA) Structure 81 

CSM Systems o}" 

LUNAR MODULE STRUCTURES, WEIGHT 84 

Ascent Stage ll~ 

Descent Stage 8b 

Lunar Module Systems-' 86-89 

AP„OLLO 13 CREW AND CREW EQUIPMENT 90 

Life Support Equipment-Space Suits 90-94 

Apollo Lunar Hand Tools 95-98 

Apollo 13 Crew Menu 99-100 

Personal Hygiene 1 01 

Medical Kit 10 ^ 

Survival Gear 102 

Biomedical Inflight Monitoring 102-10 3 

Training ^n*^?? 

Crew Biographies J - Ub lld 

LAUNCH COMPLEX 39 11 r~ 1 ^0 

MISSION CONTROL CENTER ; 116-1 IB 

MANNED SPACE FLIGHT NETWORK 119-124 

Network Computers 124-125 

Apollo Ship Vanguard 125 

Apollo Range Instrumentation Aircraft (ARIA) 125 

CONTAMINATION CONTROL PROGRAM 12 6-12 7 

LUNAR RECEIVING LABORATORY (LRL) 128-130 

LUNAR RECEIVING LABORATORY TENTATIVE SCHEDULE 131-132 

SCHEDULE FOR TRANSPORT OF SAMPLES, SPACECRAFT AND CREW 133 

APOLLO PROGRAM MANAGEMENT !34 

Apollo/Saturn Officials 13 o~ 1 ^J 

Major Apollo/Saturn V Contractors ■ 138-139 



- - 



ftj C \U C NATIONAL AERONAUTICS AND SPACE ADMINISTRATION (202) 9&2-K155 

NEW J I^VKl Washington, D.C. 20546 flELS: (202) 9*>3-6925 



FOR RELEASE: Thursday a.m. 

April 2, 1970 



RELEASE NO: 70-50 



APOLLO 13 THIRD LUNAR LANDING MISSION 



Apollo 13, the third U.S. manned lunar landing mission, 
will be launched April 11 from Kennedy Space Center, Pla., to 
explore a hilly upland region of the Moon and bring back rocks 
perhaps five billion years old. 

The Apollo 13 lunar module will stay on the Moon more 
than 33 hours and the landing crew will leave the spacecraft 
twice to emplace scientific experiments on the lunar surface 
and to continue geological investigations. The Apollo 13 
landing site is in the Pra Mauro uplands; the two National 
Aeronautics and Space Administration previous landings were in 
mare or "sea" areas, Apollo 11 in the Sea of Tranquility and 
Apollo 12 in the Ocean of Storms . 

Apollo 13 crewmen are commander James A. Lovell, Jr.; 
command module pilot Thomas K. Mattingly III, and lunar module 
pilot Fred W. Haise, Jr. Lovell is a U.S. Navy captain, 
Mattingly a Navy lieutenant commander, and Haise a civilian. 



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3/26/70 



-2- 

Launch vehicle is a Saturn V. 
Apollo 13 objectives are: 

* Perform selenologlcal Inspection, survey and 
sampling of materials in a preselected region of the 
Pra Mauro formation. 

* Deploy and activate an Apollo Lunar Surface 
Experiment Package (ALSEP) # 

* Develop man's capability to work in the lunar 
environment. 

* Obtain photographs of candidate exploration sites , 

Currently 11 television transmissions in color are 
scheduled: one In Earth orbit an hour and a half after 
launch, three on the outward voyage to the Moon; one of the 
landing site from about nine miles up; two from the lunar 
surface while the astronauts work outside the spacecraft ; 
one at the command service module/lunar module docking operation; 
one of the Moon from lunar orbit; and two on the return trip. 

The Apollo 13 landing site is in the hilly uplands to 
the north of the crater Pra Mauro. Lunar coordinates for the 
landing site are 3.6 degrees south latitude by 17.5 degrees 
west longitude, about 95.6 nautical miles east of the Apollo 
12 landing point at Surveyor III crater. 



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-3- 



Experiments emplaced at the Fra Mauro site as part of 
the ALSEP III will gather and relay long-term scientific 
data to Earth for at least a year on the Moon's physical and 
environmental properties. Five experiments are contained in 
the ALSEP: a lunar passive seismometer will measure and relay 
meteoroid impacts and moonquakes; a heat flow experiment will 
measure the heat flux from the lunar interior to the surface 
and conductivity of the surface materials to a depth of about 
10 feet; a charged particle lunar environment experiment will 
measure protons and electrons to determine the effect of the 
solar wind on the lunar environment; a cold cathode gauge 
experiment will measure density and temperature variations in 
the lunar atmosphere; and a dust detector experiment. 

The empty third stage of the Saturn V launch vehicle 
will be targeted to strike the Moon before the lunar landing 
and its impact will be recorded by the seismometer left by 
the Apollo 12 astronauts last November. The spent lunar module 
ascent stage, as in Apollo 12, will be directed to impact the 
Moon after rendezvous and final LM separation to provide a 
signal to both seismometers. 



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_4- 

Candidate future Apollo landing sites — Censorinus, 
Davy Rille, and Descartes — will be photographed with a 
large -format lunar topographic camera carried for the first 
time on Apollo 13. The lunar topographic camera will make high- 
resolution 4.5 inch square black-and-white photos in overlapping 
sequence for mosaics or in single frames. The camera mounts in 
the command module crew access hatch window when in use. After 
lunar orbit rendezvous with the lunar module and LM jettison 
the command module will make a plane -change maneuver to drive 
the orbital track over Descartes and Davy Rille for topographic 
photography. 

The Apollo 13 flight profile in general follows those 
flown by Apollos 11 and 12 with one major exception: lunar orbit 
insertion burn no. 2 has been combined with descent orbit 
insertion and the docked spacecraft will be placed into a 7x57 
nautical mile lunar orbit by use of the service propulsion 
system. Lunar module descent propellant is conserved by 
combining these maneuvers to provide 15 seconds of additional 
hover time during the landing. 

Lunar surface touchdown is scheduled to take place at 
9:55 p.m. EST April 15, and two periods of extravehicular 
activity are planned at 2:13 a.m. EST April 16 and 9:58 p.m. 
EST April 16. The LM ascent stage will lift off at 7:22 a.m. 
April 17 to rejoin the orbiting command module after more than 
33 hours on the lunar surface. 

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-5- 

Apollo 13 will leave lunar orbit at 1:42 p.m. EST April 
18 for return to Earth. Splashdown in the mid-Pacific just 
south of the Equator will be at 3:17 p.m. EST April 21 % 

After the spacecraft has landed, the crew will put on 
clean coveralls and filter masks passed in to them through 
the hatch by a swimmer, and then transfer by helicopter to a 
Mobile Quarantine Facility (MQP) on the USS Iwo Jima. The 
MQF and crew will be offloaded in Hawaii and placed aboard a 
C-lMl aircraft for the flight back to the Lunar Receiving 
Laboratory at the Manned Spacecraft Center in Houston. The 
crew will remain in quarantine up to 21 days from completion 
of the second EVA. 

The crew of Apollo 13 selected the call signs Odyssey for 
the command/service module and Aquarius for the lunar module. 
When all three crewmen are aboard the command module, the call 
sign will be "Apollo 13." As in the two previous lunar landing 
missions, an American flag will be emplaced on the lunar surface 
A plaque bearing the date of the Apollo 13 landing and the crew 
signatures is attached to the LM. 

Apollo 13 backup crewmen are USN commander John W. Young, 
commander; civilian John L. Swigert, Jr., command module pilot; 
and USAF Major Charles M. Duke, Jr., lunar module pilot. 

-more- 



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-10- 



APOLLO 13 COUNTDOWN 



Frecount activities for the Apollo 13 launch begin about 
T-4 days, when the space vehicle will be prepared for the 
start of the Official countdown. During precount, final 
space vehicle ordnance installation and electrical connections 
will be accomplished. Spacecraft gaseous oxygen and gaseous 
helium systems will be serviced, spacecraft batteries will be 
installed, and LM and CSM mechanical buildup will be completed. 
The CSM fuel cells will be activated and CSM cryogenics (liquid 
oxygen - liquid hydrogen) will be loaded and pressurised. 

The countdown for Apollo 13 will begin at T-28 hours and 
will continue to T-9 hours, at which time a built-in hold is 
planned prior to the start of launch vehicle propellant loading, 

Following are some of the major operations in the 
final count : 



T-28 hours 



Official countdown starts 

LM stowage and cabin close out 

(T-31:30 to T-18:00) 



T-27 hours, 30 minutes 



Install and connect LV flight 
batteries (to T-23 hours) 



T-22 hours, 30 minutes 



T-19 hours, 30 minutes 



T-16 hours 



Topoff of LM super critical helium 
(to T-20 hours, 30 minutes) 

LM SHe thermal shield installation 
(to T-15 hours, 30 minutes) CSM 
crew stowage (T-19 to T-12 hours, 
30 minutes) 

LV range safety checks (to T-15 
hours) 



T-15 hours 



Installation of ALSEP PCA to 
T-14 hours, 45 minutes) 



T-ll hours, 30 minutes 



Connect LV safe and arm devices 
(to 10 hours, 45 minutes) CSM pre- 
ingress operations (to T-8 hours 
45 minutes) 



T-10 hours, 15 minutes 



Start MSS move to parksite 



-more- 



-11- 



T-9 hours 



T-8 hours, 05 minutes 



T-4 hours, 17 minutes 

T-4 hours, 02 minutes 

T-3 hours, 32 minutes 

T-3 hours, 30 minutes 

T-3 hours, 07 minutes 

T-2 hours, 55 minutes 
T-2 hours, ^0 minutes 
T-2 hours 

T-l hour, 55 minutes 
T-l hour, 51 minutes 

T-43 minutes 

T-42 minutes 
T-<40 minutes 

T-30 minutes 

T-20 minutes to T-10 minutes 

T-15 minutes 



Built-in hold for 9 hours and 13 
minutes. At end of hold, pad is 
cleared for LV propellant loading 

Launch vehicle propellant loading - 
Three stages (LOX in first stage, 
LQX and LH ? in second and third 
stages). Continues thru T-3 
hours 38 minutes 

Flight crew alerted 

Medical examination 

Breakfast 

One-hour hold 

Depart Manned Spacecraft Operations 
Building for LC-39 via crew transfer 
van. 

Arrive at LC-39 

Start flight crew ingress 

Mission Control Center - Houston/ 
spacecraft command checks 

Abort advisory system checks 

Space Vehicle Emergency Detection 
System (EDS) test 

Retract Apollo access arm to stand- 
by position (12 degrees) 

Arm launch escape system 

Final launch ■ vehicle range safety 
checks (to 35 minutes) 

Launch vehicle power transfer test 
LM switch over to internal power 

Shutdown LM operational 
instrumentation 

Spacecraft to internal power 



-more- 



•12- 



T-6 minutes 

T-5 minutes. 30 seconds 

T-5 minuter* 

T-3 minutes, 7 seconds 

T-50 seconds 

T-8.9 seconds 
T-2 seconds 
T-0 



Space vehicle final status checks 

Arm destruct system 

Apollo access arm fully retracted 

Firing command (automatic 
sequence) 

Launch vehicle transfer to internal 
power 

Ignition sequence start 

All engines running 

Liftoff 



Note: Some changes in the above countdown are possible as a 
result of experience gained in the countdown demonstration test 
which occurs about 10 days before launch. 

Lightning Precautions 

During the Apollo 12 mission the space vehicle was 
subjected to two distinct electrical discharge events. 
However, no serious damage occurred and the mission pro- 
ceeded to a successful conclusion. Intensive investigation 
led to the conclusion that no hardware changes were necessary 
to protect the space vehicle from similar events. For Apollo 
13 the mission rules have been revised to reduce the pro- 
bability that the space vehicle will be launched into cloud 
formations that contain conditions conducive to initiating 
similar electrical discharges although flight into all 
clouds is not precluded. 



-more- 



-13- 



May Launch Opportunities 

The three opportunities established for May — In case 
the launch Is postponed from April 11 — provide, in effect, 
the flexibility of a choice of two launch attempts. The 
optimum May launch window occurs on May 10. The three day 
window permits a choice of attempting a launch 24 hours 
earlier than the optimum window and if necessary a further 
choice of a 24 hour or 48 hour recycle. It also permits a 
choice of making the first launch attempt on the optimum 
day with a 2 4 -hour recycle capability. The May 9 window (T-24 
hrs) requires an additional 24 hours in lunar orbit before 
initiating powered descent to arrive at the landing site at 
the same time and hence have the same Sun angle for landing 
as on May 10. Should the May 9 window launch attempt be 
scrubbed, a decision will be made at that time, based on the 
reason for the scrub, status of spacecraft cryogenics and 
weather predictions, whether to recycle for May 10 (T-0 hrs) 
or May 11 (T+24 hrs). If launched on May 11, the flight 
plan will be similar for the May 10 mission but the Sun 
elevation angle at lunar landing will be 18.5° instead of 
7.8°. 



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-15- 

LAUNCH, MISSION TRAJECTORY AND MANEUVER DESCRIPTION 

The information presented here is based on an on-time 
April 11 launch and is subject to change before or during 
the mission to meet changing conditions. 

Launch 

A Saturn V launch vehicle will lift the Apollo 13 space- 
craft from Launch Complex 39 A, NASA-Kennedy Space Center, Fla. 
The azimuth may vary from 72 to 96 degrees, depending on the 
time of launch. The azimuth changes with launch time to 
permit a fuel-optimum injection from Earth parking orbit to 
a free-return circumlunar trajectory. 

April 11 launch plans call for liftoff at 2:13 p.m. EST 
on an azimuth of 72 degrees. The vehicle will reach an altitude 
of 36 nautical miles before first stage cutoff 51 nm downrange . 
During the 2 minutes 44 seconds of powered flight, the first 
stage will increase vehicle velocity to 7,775 feet per second* 
First stage thrust will reach a maximum of 8,995 »108 pounds 
before center engine cutoff. After engine shutdown and 
separation from the second stage, the booster will fall into 
the Atlantic Ocean about 364 nm downrange from the launch site 
(30 degrees North latitude and 74 degrees West longitude) 
about 9 minutes 4 seconds after liftoff. 

The second stage (S-II) will carry the space vehicle to 
an altitude of 102 nm and a distance of 892 nm downrange. At 
engine shutdown, the vehicle will be moving at a velocity of 
21,508 fps . The four outer J-2 engines will burn 6 minutes 
32 'seconds during the powered phase, but the center engine 
will be cut off 4 minutes 47 seconds after S-II ignition. 

At outboard engine cutoff, the S-II will separate and, 
following a ballistic trajectory, plunge Into the Atlantic 
about 2,450 nm downrange from the Kennedy Space Center (31 
degrees North latitude and 33.4 degrees West longitude) some 
20 minutes 41 seconds after liftoff. 

The single engine of the Saturn V third stage (S-IVB) 
will Ignite about 3 seconds after the S-II stage separates. 
The engine will fire for 143 seconds to insert the space 
vehicle into a circular Earth parking orbit of 103 nm begin- 
ing about 1,468 nm downrange. Velocity at Earth orbital 
insertion will be 24,243 fps at 11 minutes 55 second ground 
elapsed time (GET). Inclination will be 33 degrees to the 
equator. 

*NOTE: Multiply nautical miles by 1.1508 to obtain statute 
miles; multiply feet per second by 0.6818 to obtain 
statute miles per hour. 



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-16- 



The crew will have a backup to launch vehicle guidance 
during powered flight. If the Saturn instrument unit inertial 
platform fails, the crew can switch guidance to the command 
module systems for first-stage powered flight automatic 
control. Second and third stage backup guidance is through 
manual takeover in which spacecraft commander hand controller 
inputs are fed through the command module computer to the Saturn 
instrument unit. 



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Earth Parking Orbit (EPO) 

Apollo 13 will remain in Earth parking orbit for one and 
one-half revolutions. The final "go" for the TLI burn will be 
given to the crew through the Carnarvon, Australia, Manned 
Space Flight Network station. 

Trans lunar Injection (TLI) 

Midway through the second revolution in Earth parking or- 
bit, the S-IVB third-stage engine will restart at 2:35 GET over 
the mid-Pacific Ocean near the equator and burn for almost six 
minutes to inject Apollo 13 toward the Moon. The velocity will 
increase from 25,593 fps to 36,030 fps at TLI cutoff to a free 
return circumlunar trajectory from which midcourse corrections 
could be made with the SM RCS thrusters. 

Transposition, Docking, and Ejection (TD&E) 

After the TLI burn, the Apollo 13 crew will separate the 
command/service module from the spacecraft module adapter (SLA), 
thrust out away from the S-IVB, turn around and move back in 
for docking with the lunar module. Docking should take place 
at about three hours and 21 minutes GET. After the crew con- 
firms all docking latches solidly engaged, they will connect 
the CSM-to-LM umbilicals and pressurize the LM with oxygen from 
the command module surge tank. At about 4:00 GET, the space- 
craft will be ejected from the spacecraft LM adapter by spring 
devices at the four LM landing gear "knee" attach points. The 
ejection springs will impart about one fps velocity to the 
spacecraft. A 9.4 fps S-IVB attitude thruster evasive maneuver 
in plane at 4:19 GET will separate the spacecraft to a safe dis- 
tance from the S-IVB. 

Saturn Third Stage Lunar Impact 

Through a series of pre-set and ground-commanded operations, 
the S-IVB stage /instrument unit will be directed to hit the Moon 
within a target area 375 nautical miles in diameter, centered 
Just east of Lansberg D Crater (3 degrees South latitude; 30 
degrees West longitude), approximately 124 miles west of the 
Apollo 12 landing site. 

The planned impact will provide a seismic event for the 
passive seismometer experiment placed on the lunar surface by 
the Apollo 12 astronauts in November 1969 . 

The residual propellants in the S-IVB will be used to at- 
tempt the lunar impact. Part of the remaining liquid oxygen 
(LOX) will be dumped through the engine for 48 seconds to slow 
the vehicle into a lunar impact trajectory. The liquid hydrogen 
tank's continuous venting system will vent for five minutes. 



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A mid-course correction will be made with the stage's 
auxiliary propulsion system (APS) ullage motors. A second 
APS burn will be used if necessary, at about 9 hours GET, to 
further adjust the impact point. Burn time and attitude will 
be determined from onboard systems and tracking data provided 
to ground controllers by the Manned Space Flight Network. 

The LOX dump by itself would provide a lunar impact; the 
mid-course correction burns will place the S-IVB/IU within the 
desired target area for impact about 20 minutes after the com- 
mand/service module enters lunar orbit. 



Time 
Hrs :Min 



The schedule of events concerning the lunar impact is: 

Event 



02 42 

04 19 

04 21 

04 3& 

04 41 

04 41 

04 42 

06 00 

08 59 

09 04 
77 46 

Translunar Coast 



Translunar injection (TLI) — 
maneuver completion 

Begin S-IVB evasive maneuver (APS 
engines ) 

End evasive maneuver 

LH ? tank continuous vent on 

Begin LOX dump 

LHp tank continuous vent off 

End LOX dump 

Begin first APS burn 

Begin final APS burn (if required) 

APS ullage engines off 

Lunar impact of S-IVB/IU 



Up to four midcourse correction burns are planned during 
the spacecraft's translunar coast, depending upon the accuracy 
of the trajectory resulting from the TLI maneuver. If required, 
the midcourse correction burns are planned at TLI+9 hours, TLI+ 
30 hours, 41 minutes, lunar orbit insertion (LOI)-22 hours and 
LOI-5 hours. The MCC-2 is a 15 fps SPS hybrid transfer maneuver 
which lowers pericynthion from 210 nm to 59 nm and places Apollo 
13 on a non-free-return trajectory. 



■more- 



-27- 



Return to the free-return trajectory is always within 
the capability of the spacecraft service propulsion or des- 
cent propulsion systems. 

During coast periods between midcourse corrections, the 
spacecraft will be in the passive thermal control (PTC) or 
"barbecue" mode in which the spacecraft will rotate slowly 
about its roll axis to stabilize spacecraft thermal response 
to the continuous solar exposure. 

Lunar Orbit Insertion (LOT) 

The lunar orbit insertion burn will be made at 77:25 GET 
at an altitude of about 85 nm above the Moon. The LOI burn 
will have a nominal retrograde velocity change of 2815 fps and 
will insert Apollo 13 into a 57x168 nm elliptical lunar orbit. 

Descent Orbit Insertion (DPI) 

A 213 fps SPS retrograde burn at 81:45 GET will place the 
CSM /LM into a 7x57 nm lunar orbit from which the LM will begin 
the later powered descent to landing. In Apollos 11 and 12, 
DOI was a separate maneuver using the LM descent engine. The 
Apollo 13 DOI maneuver in effect is a combination LOI-2 and DOI 
and produces two benefits: conserves LM descent propellant 
that would have been used for DOI and makes this propellant 
available for additional hover time near the surface, and allows 
11 lunar revolutions of spacecraft tracking in the descent orbit 
to enhance position/velocity (state vector) data for updating 
the LM guidance computer during the descent and landing phase. 

Lunar Module Separation 

The lunar module will be manned and checked out for un- 
docking and subsequent landing on the lunar surface north of 
the crater, Pra Mauro* Undocking during the 12th revolution 
will take place at 99:16 GET. A radially downward service 
module RCS burn of 1 fps will place the CSM on an equiperiod 
orbit with a maximum separation of 2.5 nm. 

CSM Circularization 

During the 12th revolution, a 70 fps posigrade SPS burn 
at 100:35 GET will place the CSM into 52x62 nm lunar orbit, 
which because of perturbations of the lunar gravitational po- 
tential, should become nearly circular at the time of rendez- 
vous with the LM. 



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Power Descent Initiation (,PDI), Lunar Landing 

During the 14th revolution a three-phase powered descent 
(PD) maneuver begins at pericynthion at 103:31 GET using the 
LM descent engine to brake the vehicle out of the descent orbit. 
The guidance-controlled PD maneuver starts about 260 nm prior 
to touchdown, and is in retrograde attitude to reduce velocity 
to essentially zero at the time vertical descent begins. Space- 
craft attitude will be windows up from powered descent initia- 
tion to the end of the braking phase so that the LM landing 
radar data can be integrated continually by the LM guidance 
computer and better communications can be maintained. The 
braking phase ends at about 7,400 feet above the surface and 
the spacecraft is rotated more toward an upright windows-forward 
attitude to permit a view of the landing site. The start of the 
approach phase is called high gate, and the start of the landing 
phase at about 500 feet is called low gate. 

Both the approach (visibility) phase and landing phase 
allow pilot takeover from guidance control as well as visual 
evaluation of the landing site. The final vertical descent to 
touchdown begins at about 100 feet when all forward velocity is 
nulled out. Vertical descent rate will be 3 f ps . The crew may 
elect to take over manual control at approximately 500 feet. 
The crew will be able to return to automatic landing control 
after a period of manned maneuvering if desirable. Touchdown 
will take place at 103:42 GET. 

Lunar Surface Exploration 

During the 33 1/2 hours Apollo 13 commander James Lovell 
and lunar module pilot Fred Haise are on the surface, they will 
leave the lunar module twice for four-hour EVAs . These are ex- 
tendable to five hours in real time if the physical conditions 
of the astronauts and amount of remaining consumables permit. 

In addition to gathering more data on the lunar environ- 
ment and bringing back geological samples from a third lunar 
landing site, Lovell and Haise will deploy a series of experi- 
ments which will relay back to Earth long-term scientific 
measurements of the Moon's physical and environmental properties 

The experiments series, called the Apollo Lunar Surface 
Experiment Package (ALSEP), will be left on the surface and 
could transmit scientific and engineering data to the Manned 
Space Plight Network for at least a year. 

The ALSEP for Apollo 13, stowed in the LM descent stage 
scientific equipment bay, comprises components for the five 
ALSEP experiments — passive seismic, heat flow, charged par- 
ticle lunar environment, cold cathode gauge, and lunar dust 
detector . 

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These experiments are aimed toward determining the 
structure and state of the lunar interior, the composition 
and structure of the lunar surface and processes which modify 
the surface, and evolutionary sequence leading to the Moon's 
present characteristics. The Passive Seismic Experiment will 
become the second point in a lunar seismic net begun with the 
first ALSEP at the Surveyor III landing site of Apollo 12. 
Those two seismometers must continue to operate until the next 
seismometer is emplaced to complete the three-station set. The 
heat flow experiment includes drilling two 10-foot holes with 
the lunar surface drill. 

While on the surface, the crew's operating radius will be 
limited by the range provided by the oxygen purge system (OPS), 
the reserve oackup for each man's portable life support system 
(PLSS) backpack. The OPS supplies ^5 minutes of emergency 
breathing oxygen and suit pressure. 

Among other tasks assigned to Lovell and Haise for the 
two EVA periods are: 

^Collect a contingency sample of about two pounds of 
lunar material. 

^Gather about 95 pounds of representative lunar surface 
material, including core samples, individual rock samples and 
fine-grained fragments from the Fra Mauro hilly uplands site. 
The crew will photograph thoroughly the areas from which sam- 
ples are taken. 

*Make observations and gather data on the mechanical 
properties and terrain characteristics of the lunar surface 
and conducting other lunar field geological surveys, including 
digging a two-foot deep trench for a soil mechanics investigation 

*Photograph with ";he lunar stereo closeup camera small 
geological features that would be destroyed in any attempts to 
gather them for return to Earth. 

^Deploy and retrieve a windowshade-like solar wind compo- 
sition experiment similar to the ones used in Apollos 11 and 12. 

Early in the first EVA, Lovell and Haise will set up the 
erec table S-Band antenna near the LM for relaying voice, TV, 
and LM telemetry to MSFN stations. After the antenna is de- 
ployed, Haise will climb back into the LM to switch from the 
LM steerable S-Band antenna to the erectable antenna while 
Lovell makes final adjustments to the antenna's alignment. 
Haise will then rejoin Lovell on the lunar surface to set up 
a United States flag and continue with EVA tasks. 



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Red stripes around the elbows and knees of Lovell's 
pressure suit will permit crew recognition during EVA tele- 
vision transmissions and on photographs. 

Ascent, Lunar Orbit Rendezvous 

Following the 33-hour lunar stay the LM ascent stage will 
lift off the lunar surface to begin the rendezvous sequence ■ . 
with the orbiting CSM. Ignition of the LM ascent engine will 
be at 137:09 for a seven minute eight second burn attaining a 
total velocity of 6,044 fps. Powered ascent is in two phases: 
vertical ascent for terrain clearance and the orbital inser- 
tion phase. Pitchover along the desired launch azimuth begins 
as the vertical ascent rate reaches 50 fps about 10 seconds 
after liftoff at about 272 feet in altitude. Insertion into a 
9x44 nm lunar orbit will take place about 166 nm west of the 
landing site. 

Following LM insertion into lunar orbit, the LM crew will 
compute onboard the major maneuvers for rendezvous with the CSM 
which is about 267 nm ahead of and 51 miles above the LM at this 
point. All maneuvers in the sequences will be made with the LM 
RCS thrusters. The premission rendezvous sequence maneuvers, 
time, and velocities, which likely will differ slightly in real 
time, are as follows: 

Concentric sequence Initiate (CSI) : At first LM apolune 
after insertion, 138:19 GET, 50 fps posigrade, following some 
20 minutes of LM rendezvous radar tracking and CSM sextant/VHP 
ranging navigation. CSI will be targeted to place the LM in 
an orbit 15 nm below the CSM at the time of the later constant 
delta height (CDH) maneuver (139:04). 

The CSI burn may also initiate corrections for any out-of- 
plane dispersions resulting from insertion azimuth errors. The 
resulting LM orbit after CSI will be 45x43.5 nm and will have a 
catchup rate to the CSM of about 120 feet per second. 

Terminal phase initiation (TPI): This maneuver occurs at 
139:46 and adds 24.7 fP s along the line of sight toward the 
CSM when the elevation angle to the CSM reaches 26.6 degrees. 
The LM orbit becomes 61x44 nm and the catchup rate to the CSM 
decreases to a closing rate of 133 fps. 

Midcourse correction maneuvers will be made if needed, fol- 
lowed by four braking maneuvers. Docking nominally will take 
place at 140:25 GET to end the three and one-half hour rendez- 
vous sequence. 

The LM ascent stage will be jettisoned atl43:04 GET and 
a CSM RCS 1.0 fps maneuver will provide separation. 



-more- 



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Ascent Stage Deorbit 

Prior to transferring to the command module , the LM crew 
will set up the LM guidance system to maintain the ascent stage 
in an inertial attitude. At about 1*44:32 GET the LM RCS thrus- 
ters will ignite on ground command for 186 fps retrograde burn 
targeted for ascent stage impact at 145:00 about 35 miles from 
the landing site. The burn will have a small out-of-plane north 
component so that the ground track will include the original 
landing site. The ascent stage will impact at about 5508 fps 
at an angle of four degrees relative to the local horizontal. 
The ascent stage deorbit serves to remove debris from lunar or- 
bit. Impacting an object with a known velocity and mass near 
the landing site will provide experimenters with an event for 
calibrating readouts from the ALSEP seismometer left behind. 

A plane change maneuver at 154:13 GET will place the CSM 
on an orbital track passing directly over the crater Descartes 
and Davy Rille eight revolutions later for photographs from 
orbit. The maneuver will be a 825 fps/SPS burn out of plane 
for a plane change of 8.8 degrees, and will result in an orbit 
inclination of 11.4 degrees. 

Transearth Injection (TEI) 

The nominal transearth injection burn will be at 167:29 
GET following 90 hours in lunar orbit. TEI will take place on 
the lunar farside, will be a 3,147 fps posigrade SPS burn of 
two minutes 15 seconds duration and will produce an entry 
velocity of 36,129 fps after a 72 hours transearth flight time. 

Transearth Coast 



Three entry corridor-control transearth midcourse correc- 
tion burns will be made if needed: MCC-5 at TEI+15 hours, 
MCC-6 at entry interface (EI) -22 hours and MCC-7 at EI -3 hrs . 

Entry , Landing 

Apollo 13 will encounter the Earth's atmosphere (400,000 
feet) at 240:50 GET at a velocity of 36,129 fps and will land 
approximately 1,250 nm downrange from the entry-interface point 
using the spacecraft's lifting characteristics to reach the 
landing point. Splashdown will be at 241:04 at 1.5 degrees 
South latitude by 157*5 degrees West longitude. 



-more- 



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Recovery Operations 

Launch abort landing areas extend downrange 3,400 nautical 
miles from Kennedy Space Center, fanwise 50 nm miles above and 
below the limits of the variable launch azimuth (72-96 degrees) 
in the Atlantic Ocean. On station in the launch abort area will 
be the destroyer USS Mew. 

The landing platform-helicopter (LPH) Iwo Jima, Apollo 13 
prime recovery ship, will be stationed near the Pacific Ocean 
end-of -mission aiming point prior entry. 

Splashdown for a full-duration lunar landing mission 
launched on time April 11 will be at one degree 34 minutes 
South by 157 degrees 30 minutes West about 180 nautical miles 
South of Christmas Island, at 24l:04 GET (3:17 p.m. EST) April 
21. 

In addition to the primary recovery vessel located on the 
mid-Pacific recovery line and the surface vessel in the launch 
abort area, eight HC-130 aircraft will be on standby at five 
staging bases around the Earth: Guam; Hawaii; Azores; 
Ascension Island;and Florida. 

Apollo 13 recovery operations will be directed from the 
Recovery Operations Control Room in the Mission Control Center, 
supported by the Atlantic Recovery Control Center, Norfolk, Va., 
and the Pacific Recovery Control Center, Kunia, Hawaii, 

After splashdown, the Apollo 13 crew will don clean cover- 
alls and filter masks passed to them through the spacecraft 
hatch by a recovery swimmer. The crew will be carried by heli- 
copter to the Iwo Jima where they will enter a Mobile Quaran- 
tine Facility (MQF) about 90 minutes after landing. 



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APOLLO 13 ONBOARD TELEVISION 

Apollo 13 will carry two color and one black-and-white 
television cameras. One color camera will be used for 
command module cabin interiors and out-the-window Earth/ 
Moon telecasts, and the other color camera will be stowed 
in the LM descent stage from where it will view the astronaut 
initiate egress to the lunar surface and later will be de- 
ployed on a tripod to transmit a real-time Dicture of the two 
periods of lunar surface EVA. The black-and-white camera 
will be carried in the LM cabin. It will only be used as a 
backup \;o the lunar surface color camera. 

The two color TV cameras are essentially identical, 
except :7or additional thermal protection on the lunar surface 
camera. Built by Westinghouse Electric Corp., Aerospace 
£*P-Sf on .- Baltimore, Md. , the color cameras output a standard 
b25-lme 9 30 frane-per-second signal in color by use of a 
rotating color wheel system. 

The color TV cameras weigh 12 pounds and are fitted with 
zoom lenses for wideangle or closeup fields of view. The 
CM camera Is fitted with a three-inch monitor for framing and 
focusing. The lunar surface color camera has 100 feet of cable 
available . 

The bc.ckup black-and-white lunar surface TV camera; also 
built by Westinghouse, is of the same type used in the first 
manned um,r landing in Apollo 11. It weighs 7.25 pounds and 
draws 6.5 watts of 24-32 volts DC power. Scan rate is 10 
frames-per.-second at 325 lines-per-frame . The camera body is 
10.6 inches long, 6.5 inches wide and' 3-4 inches deep, and is 
fitted with bayonet-mount wideangle and lunar day lenses . 

During the two lunar surface EVA periods, Apollo 13 
commander Love 11 will be recognizable by red stripes around 
the elbows and knees of his pressure suit. 

The following is a preliminary plan for TV transmissions 
based upon a 2:13 p.m. EST April 11 launch. 



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-48- 

APOLLO 13 SCIENCE 



Lunar Orbital Photography 

Science experiments and photographic tasks will be 
conducted from the CSM during the Apollo 13 mission. During 
the translunar phase of the mission, photography will be 
taken of the Earth as well as various operational photography. 

During lunar orbit, various lunar surface features 
including candidate landing sites Censorinus, Descartes and 
Davy Rille and the Apollo 11 and 12 landing sites will be 
photographed with the Lunar Topographic Camera. In addition, 
five astronomical phenomena will be photographed: 

1) Photographs will be taken of the solar corona using 
the Moon as an occulting edge to block out the solar 
disk. 

2) Photography will be taken of the zodiacal light which 
is believed to originate from reflected sunlight in 
the astoroid belt. Earth observation of zodiacal 
light is inconclusive due to atmospheric distortion. 

3) Photography will be taken of lunar limb brightening, 
which appears as bright rim light above the horizon 
following lunar sunset. 

4) Photographs will be taken of the Comet J,C. Bennett, 
1969i which should be visible from lunar orbit during 
the Apollo 13 mission. 

5) Photographs will be taken of the region of Gegenschein 
which is a faint light source covering a 20° field of 
view about the Earth-Sun line on the opposite side of 
the Earth from the Sun (anti-solar axis). One of the 
theories for the Gegenschein source is the existence 
of trapped particles of matter at the Moulton point 
which produce brightness due to reflected sunlight. 
The Moulton point is a theoretical point located 
940,000 statute miles from the Earth along the anti- 
solar axis at which the sum of all gravitational forces 
is zero. Prom the vantage point of lunar orbit, the 
Moulton point region may be photographed from approxi- 
mately 15° off the Earth/Sun line. These photographs 
should show if Gegenschein results from the Moulton 
point theory or from zodiacal light or a similar source 



■more- 



-H9- 



Photographic studies will be made on Apollo 13 of the 
ice particle flow following a water dump and of the gaseous 
cloud which surrounds a manned spacecraft in a vacuum and 
results from liquid dumps, outgassing, etc. 

In addition to the photographic studies, an experiment 
will be conducted with the CSM VHP communications link. 
During this experiment, the VHP signal will be reflected from 
the lunar surface and received by a 150-foot antenna on Earth. 
By analysis of the wavelength of the received signal, certain 
lunar subsurface characteristics may be discernible such as 
the depth of the lunar regolith layer. This experiment is 
called VHF Bistatic Radar. 

Charged Particle Lunar Environment Experiment (CPLEE) 

The scientific objective of the Charged Particle Lunar 
Environment Experiment is to measure the particle energies of 
protons and electrons that reach the lunar surface from the 
Sun. Increased knowledge on the energy distribution of these 
particles will help us understand how they perturb the Earth- 
Moon system. At some point electrons and protons in the 
magnetospheric tail of the Earth are accelerated and plunge 
into the terrestrial atmosphere causing the spectacular 
auroras and the Van Allen radiation. When the Moon is in 
interplanetary space the CPLEE measures proton and electrons 
from solar flares which results in magnetic storms in the 
Earth's atmosphere. Similar instruments have been flown on 
Javelin rockets and on satellites. The lunar surface, however, 
allows data to be gathered over a long period of time and from 
a relatively stable platform in space. 

To study these phenomena, the CPLEE measures the energy 
of protons and electrons simultaneously from 50 electron 
volts to 50,000 electron volts (50Kev)_. The solar radiation 
phenomena measured are as follows: 

a. Solar wind electrons and protons 50ev-5Kev. 

b. Thermalized solar wind protons and electrons 
50ev-10Kev. 

c. Magnetospheric tail particles 50ev to 50Kev. 

d. Low energy solar cosmic rays 40ev-50Kev. 



-more- 



iiiH 



-50- 



PARTICLES IN 




COLUMATING 
SLITS 



CPLEE PHYSICAL ANALYZER 




PHYSICAL ANALYZER 
ELECTRONICS 



CHARGED-PARTICLE 

LUNAR ENVIRONMENT 

EXPERIMENT SUBSYSTEM 



-more- 



-51- 

Thls experiment is distinct from the ALSEP Solar Wind 
Spectrometer (SWS) flown on Apollo 12 which measures 
direction as well as energy levels. The SWS measures elec- 
trons from 10 . 5ev to l,400ev and protons from 75ev to 
10,000ev. 

The detector package contains two spectrometers providing 
data on the direction of the incoming flux. 

Each spectrometer has six particle detectors: five C- 
shaped channeltron photon-multipliers and one funneltron, a 
helical shaped photon multiplier. Particles of a given charge 
and different energies on entering the spectrometer are subject 
to varying voltages and deflected toward the five channeltrons 
while particles of the opposite charge are deflected toward 
the funneltron. Thus electrons and protons are measured 
simultaneously in six different energy levels . The voltages 
are changed over six steps; +35V, +350 volts and +3500V. In 
this way electrons and protons are measured from 50ev to 70Kev 
in a period of less than 20 seconds. 

The channeltron is a glass capillary tube having an inside 
diameter of about one millimeter and a length of 10 centimeters. 
The helical funneltron has an opening of 8mm. When a voltage 
is applied between the ends of the tube, an electric field is 
established down its length. Charged particles entering the 
tube are amplified by a factor of -, 8. 

The spectrometers have two ranges of sensitivity and can 

4 10 2 

measure fluxes between 10 and 10 particles/cm -sec-steradian. 

The charged particle lunar environment experiment (CPLEE) 
and data analysis are the responsibility of Dr. Brian O'Brien, 
University of Sydney (Australia) and Dr. David Reasoner, Rice 
University, with Dr. O'Brien assuming the role of Principal 
Investigator. 

Lunar Atmosphere Detector (LAD) 

Although the Moon is commonly described as a planetary 
body with no atmosphere, the existence of some atmosphere cannot 
be doubted. Two sources of this atmosphere are predicted: 
internal, i.e., degassing from the interior of the Moon either 
by constant diffusion through its surface or intermittent 
release from active vents; external i.e., solar wind and 
vaporization during meteorite impacts. Telescopic observations 
from polarized scattered light indicate that the atmospheric 
pressure could not exceed one millionth of a torr (a torr is 
defined as 1/760 of the standard atmosphere). 



-more- 



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Measurements will be of the greatest significance if 
it turns out through later orbital sensors that they are of 
internal origin. The Earth's atmosphere and oceans have 
been released from the Earth's interior by degassing. The 
most certain source, however, is the solar wind whose ionized 
particles become neutralized in the lunar atmosphere and 
then are released as neutral gases. Neon is the predominant 
gas expected. Lighter gases such as hydrogen and helium escape 
and heavier ones statistically should be present in small 
quantities. Neutral particles are ionized in the lunar atmos- 
phere, further reducing the numbers present; others will escape 
as the temperature rises (and concentrate near the surface 
when it falls) . 

The LAD utilizes a cold cathode ionization gauge to 
measure the density of neutral particles at the lunar surface 
and the variations in density association with lunar phase 
or solar activity. The ionization gauge is basically a crossed 
electro-magnetic field device. Electrons in the gauge are 
accelerated by the combined magnetic and electric fields pro- 
ducing a collision are collected by the cathode where they 
form a flow of positive ions. The positive ions current is 
found to be proportional to the density of the gas molecules 
entering the gauge. In addition, the gauge temperature is 
read over the range of -90° to 125°C with + 5°C accuracy. 

From the density and temperature data the pressure of the 
ambient lunar atmosphere can then be calculated. Chemical 
composition of the atmosphere however is not directly measured 
but the gauge has been calibrated for each gas it is expected 
to encounter on the lunar surface and some estimates can be 
made of the chemical composition. Any one of seven different 
dynamic ranges may be selected permitting detection of neutral 

—6 —12 

particles from 10"" Torr (highest pressure predicted) to 10 

Torr (maximum capability of gauge) . For pressure greater than 

10 Torr accuracies of - 30% will be obtained; for pressures 

-10 + 

less then 10 Torr accuracies - 50? will be obtained. The 

experiment, therefore, will reduce the present uncertainty from 

a magnitude to a factor. 

The Lunar Atmosphere Detector (LAD) and data are the 
responsibility of Francis Johnson, University of Texas (Dallas) 
and Dallas Evans, Manned Spacecraft Center, with Dr. Johnson 
serving as Principal Investigator. 



-more- 



-54- 

Lunar Heat Flow Experiment (HFE) 

The scientific objective of the Heat Flow experiment 
is to measure the steady-state heat flow from the lunar 
interior. Two predicted sources of heat are: 1) original 
heat at the time of the Moon's formation and 2) radioactivity. 
Scientists believe that heat could have been generated by the 
infalling of material and its subsequent compaction as the 
Moon was formed. Moreover , varying amounts of the radioactive 
elements uranium, thorium and potassium were found present 
in the Apollo 11 and 12 lunar samples which if present at 
depth, would supply significant amounts of heat. No simple 
way has been devised for relating the contribution of each of 
these sources to the present rate of heat loss. In addition 
to temperature, the experiment is capable of measuring the 
thermal conductivity of the lunar rock material* 

The combined measurement of temperature and thermal 
conductivity gives the net heat flux from the lunar interior 
through the lunar surface. Similar measurements on Earth 
have contributed basic information to our understanding of 
volcanoes, earthquakes and mountain building processes. In 
conjunction with the seismic and magnetic data obtained on 
other lunar experiments the values derived from the heat flow 
measurements will help scientists to build more exact models 
of the Moon and thereby give us a better understanding of its 
origin and history. 

The Heat Flow experiment consists of instrument probes, 
electronics and emplacement tool and the lunar surface drill. 
Each of two probes is connected by a cable to an electronics 
box which rests on the lunar surface. The electronics, which 
provide control, monitoring and data processing for the 
experiment, is connected to the ALSEP central station. 

Each probe consists of two identical 20-inch (50 cm) long 
sections each of which contains a "gradient" sensor 
bridge, a "ring" sensor bridge and two heaters. Each bridge 
consists of four platinum resistors mounted in a thin-walled 
fiberglass cylindrical shell. Adjacent areas of the bridge 
are located in sensors at opposite ends of the 20-inch fiber- 
glass probe sheath. Gradient bridges consequently measure the 
temperature difference between two sensor locations. 



-more- 



-55- 




APOLLO LUNAR SURFACE DRILL 



-more- 



-56- 



PROBE PACKAGE 
CABLE TRAY 



ELECTRONICS 
PACKAGE 



PROBE CARRYING PACKAGE 
(CONTAINS 2 PROBES & 
EMPLACEMENT TOOL) 



SUNSHIELD 



THERMAL 
MASK 




REFLECTOR 



CABLE BRACKET 
REMOVED DURING 
DEPLOYMENT 



LUNAR 

SURFACE 



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RADIATION 
SHIELD 





RADIATION 
SHIELD 




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THERMOCOUPLES 
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> PROBE 



HEAT FLOW EXPERIMENT 



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-57- 



In thermal conductivity measurements at very low values 
a heater surrounding the gradient sensor is energized with 
0.002 watt? and the gradient sensor values monitored. The 
rise in temperature of the gradient sensor is a function of 
the thermal conductivity of the surrounding lunar material. 
For higher range of values, the heater is energized at 0,5 
watts of heat and monitored by a ring sensor. The rate of 
temperature rise, monitored by the ring sensor is a function 
of the thermal conductivity of the surrounding lunar material. 
The ring sensor, approximately four inches from the heater, is 
also a platinum resistor. A total of eight thermal conduc- 
tivity measurements can be made. The thermal conductivity 
mode of the experiment will be implemented about twenty days 
(500 hours) after- deployment. This is to allow sufficient 
time for the perturbing effects, of drilling and emplacing the 
probe in the borehole to decay; i.e., for the probe and casings 
to come to equilibrium with the lunar subsurface. 

A 30-foot (10 meter) cable connects each probe to the 
electronics box. In the upper six feet of the borehole the 
cable contains four evenly spaced thermocouples: at the top 
of the probe; at 26" (65 cm), 45" (115 cm), and 66" (165 cm). 
The thermocopules will measure temperature transients pro- 
pagating downward from the lunar surface. The reference junction 
temperature for each thermocouple is located in the electronics 
box. In fact, the feasibility of making a heat flow measure- 
ment depends to a large degree on the low thermal conductivity 
of the lunar surface layer, the regolith. Measurement of lunar 
surface temperature variations by Earth-based telescopes as 
well as the Surveyor and Apollo missions show a remarkably 
rapid rate of cooling. The wide fluctuations in temperature 
of the lunar surface (from -250°F to +250°) are expected to 
influence only the upper six feet and not the bottom 3 feet 
of the borehole. 

The astronauts will use the Apollo Lunar Surface Drill 
(ALSD) to make a lined borehole in the lunar surface for the 
probes. The drilling energy will be provided by a battery- 
powered rotary percussive power head. The drill rod consists 
of fiberglass tubular sections reinforced with boron filaments 
(each about 20 inches or 50 cm long). A closed drill bit, 
placed on the first drill rod, is capable of penetrating the 
variety of rock including three feet of vesicular basalt 
(40 per cent porosity). As lunar surface penetration pro- 
gresses, additional drill rod sections will be connected to the 
drill string. The drill string is left in place to serve 
as a hole casing. 



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An emplacement tool Is used, by the astronaut to insert 
the probe to full depth. Alignment springs position the 
probe within the casing and assure a well-defined radiative 
thermal coupling between the probe and the borehole. 
Radiation shields on the hole prevent direct sunlight from 
reaching the bottom of the hole. 

The astronaut will drill a third hole near the HPE 
and obtain cores of lunar material for subsequent analysis of 
thermal properties. 

Heat flow experiment, design and data analysis are the 
responsibility of Dr. Marcus Langseth of the Lamont-Doherty 
Geological Observatory; Dr. Sydney Clark, Jr., Yale University, 
and Dr. M. G. Simmons, MIT] with Dr. Langseth assuming the 
role of Principal Investigator. 

Passive Seismic Experiment (PSE) 

The ALSEP Passive Seismic Experiment (PSE) will measure 
seismic activity of the Moon and obtain information on the 
physical properties of the lunar crust and interior. The 
PSE will detect surface tilt produced by tidal deformations, 
moonquakes and meteorite impacts. 

The passive seismometer design and subsequent experiment 
analysis are the responsibility of Dr . Gary Latham of the Lamont- 
Doherty Geological Observatory. 

A similar passive seismic experiment was deployed as part 
of the Apollo 12 ALSEP station at Surveyor crater last 
November and has transmitted Earthward lunar surface seismic 
activities since that time. The Apollo 12 and 13 seismometers 
differ from the seismometer left at Tranquility Base in July 
1969 by the Apollo 11 crew in that they are continuously 
powered by a SNAP-27 radioisotope electric generator, while 
the Apollo 11 seismometer was powered by solar energy and could 
output data only during the lunar day at its location. 

After Lovell and Haise ascend from the lunar surface and 
rendezvous with the command module in lunar orbit, the lunar 
module ascent stage will be jettisoned and later ground- 
commanded to impact on the lunar surface about 42 statute 
miles from the Apollo 13 landing site at Fra Mauro. Impact 
of an object of known mass and velocity will assist in cali- 
brating the Apollo 13 seismometer readouts as well as providing 
comparative readings between the Apollo 12 and 13 seismometers 
forming the first two stations of a lunar surface seismic net- 
work. 

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There are three major physical components of the PSE: 

* The sensor assembly consists of three, long-period 
seismometers with orthogonally-oriented, capaci- 
tance type seismic sensors, measuring along two 
horizontal axes and one vertical axis. This is 
mounted on a gimbal platform assembly. There is 
one short period seismometer which has magnet-type 
sensors. It is located directly on the base of the 
sensor assembly. 

* The leveling stool allows manual leveling of the sen- 
sor assembly by the astronaut to within +5°, and 
final leveling to within 3 arc seconds by control 
motors . 

* The thermal shroud covers and helps stabilize the 
temperature of the sensor assembly. Also, two radio- 
isotope heaters will protect the instrument from 

the extreme cold of the lunar night. 

Solar Wind Composition Experiment (SWCE) 

The scientific objective of the solar wind composition 
experiment is to determine the elemental and isotopic com- 
position of the noble gases in the solar wind. (This is 
not an ALSEP experiment). 

The solar wind composition detector experiment design 
and subsequent data analysis are the responsibility of 
J. Geiss and P. Eberhardt, University of Bern (Switzerland) 
and P. Signer, Swiss Federal Institute of Technology, with 
Professor Geiss assuming the responsibility of Principal 
Investigator . 

As in Apollo 11 and 12 the SWC detector will be deployed 
on the Moon and brought back to Earth by the astronauts. 
The detector, however, will be exposed to the solar wind flux 
for 20 hours instead of two hours as in Apollo 11 and 18 hours 
42 minutes on Apollo 12. 

The solar wind composition detector consists of an aluminum 
foil four square feet in area and about 0.5 mils thick rimmed 
by Teflon for resistance to tear during deployment. A staff 
and yard arrangement will be used to deploy the foil and to 
maintain the foil approximately perpendicular to the solar 
wind flux. Solar wind particles will penetrate into the 
foil while cosmic rays will pass right through. The solar wind 
particles will be firmly trapped at a depth of several hundred 
atomic layers. After exposure on the lunar surface, the foil 
is reeled and returned to Earth. 

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SOLAR WIND EXPERIMENT 



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Dust Detector 

The ALSEP Dust Detector Is an engineering measurement 
designed to detect the presence of dust or debris that may 
impinge on the ALSEP or accumulate during its operating 
life. 

The measurement apparatus consists of three calibrated 
solar cells, one pointing in east, west and vertical to face 
the eliptic path of the Sun, The detector is located on the 
central station. 

Dust accumulation on the surface of the three solar cells 
will reduce the solar illumination detected by the cells. 
The temperature of each cell will be measured and compared 
with predicted values. 



SOLAR CELLS 




DUST DETECTOR 
SENSOR PACKAGE 



« i CA B LE 



DUS7 DETECTOR 



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Field Geology Investigations 

The scientific objectives of the Apollo Field Geology 
Investigations are to determine the composition of the Moon 
and the processes which shape Its surfaces. This information 
will help to determine the history of the Moon and its rela- 
tionship to the Earth. Apollo 11 visited the Sea of Tran- 
quility (Mare Tranquillitatis ) and Apollo 12 studied the Ocean 
of Storms (Oceanus Procellarum) . The results of these studies 
should help establish the nature of Mare-type areas. Apollo 13 
will Investigate a hilly upland area. 

Geology investigation of the Moon actually began with the 
telescope. Systematic geology mapping began 10 years ago with 
a team of scientists at the U.S. Geological Survey. Ranger, 
Surveyor, and especially Lunar Orbit er data enormously increased 
the detail and accuracy of these studies. The Apollo 11 and 12 
investigations represent another enormous advancement in pro- 
viding new evidence on the Moon f s great age, its curious chemi- 
stry, the surprisingly high density of the lunar surface 
material. 

On Apollo 13, almost the entire second EVA will be devoted 
to the Field Geology Investigations and the collection of docu- 
mental samples. The sample locations will be carefully photo- 
graphed before and after sampling. The astronauts will care- 
fully describe the setting from which the sample is collected. 
In addition to specific tasks, the astronauts will be free to 
photograph and sample phenomena which they judge to be unusual, 
significant, and interesting. The astronauts are provided with 
a package of detailed photo maps which they will use for plan- 
ning traverses. Photographs will be taken from the LM window. 
Each feature or family of features will be described, relating 
to features on the photo maps. Areas and features where photo- 
graphs should be taken and representative samples collected will 
be marked on the maps . The crew and their ground support per- 
sonnel will consider real-time deviation from the nominal plan 
based upon an on-the-spot analysis of the actual situation. A 
trench will be dug- for soil mechanics investigations. 

The Earth-based geologists will be available to advise the 
astronauts in real-time and will work with the data returned, 
the photos, the samples of rock and the astronauts' observations 
to reconstruct here on Earth the astronauts traverse on the Moon 

Each astronaut will carry a Lunar Surface Camera (a modi- 
fied 70 mm electric Hasselblad). The camera has a 60 mm lens 
and a Reseau plate. Lens apertures range from 5/5.6 to f/^5- 
Its focus range is from three feet to infinity, A removable 
polarizing filter is attached to the lens of one of the cameras 
and can be rotated in 45-degree Increments for light polarizing 
studies . 

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A gnomon, used for metric control of near field (less 
than 10 feet) stereoscopic photography, will provide angular 
orientation relative to the local vertical. Information on 
the distances to objects and on the pitch, roll, and azimuth 
of the camera's optic axis are thereby included in each photo- 
graph. The gnomon is a weighted tube suspended vertically on 
a tripod supported gimbal. The tube extends one foot above 
the gimbal and is painted with a gray scale in bands one centi- 
meter wide. Photogrammetric techniques will be used to pro- 
duce three-dimensional models and maps of the lunar surface 
from the angular and distance relationship between specific 
objects recorded on the film. 

The 16 mm Data Acquisition Camera will provide times: se- 
quence coverage from within the LM. It can be operated in 
several automatic modes, ranging from one frame/second to 24 
frames/second. Shutter speeds, which are independent of the 
frame rates, range from 1/1000 second to 1/60 second. Time 
exposures are also possible. While a variety of lenses is 
provided, the 18 mm lens will be used to record most of the 
geological activities in the one frame/second mode. A similar 
battery powered 16 mm camera will be carried in EVA. 

The Lunar Surface Close-up Camera will be used to obtain 
very high resolution close-up stereoscopic photographs of the 
lunar surface to provide fine scale information on lunar soil 
and rock textures. Up to 100 stereo pairs can be exposed on 
the preloaded roll of 35 mm color film. The handle grip en- 
ables the astronaut to operate the camera from a standing posi- 
tion. The film drive and electronic flash are battery-operated, 
The camera photographs a 3 M x3 M area of the lunar surface. 

Geological sampling equipment includes tongs, scoop, ham- 
mer, and core tubes, A 24-inch extension handle is provided 
for several of the tools to aid the astronaut in using them 
without kneeling. 

Sample return containers (SRC) have been provided for re- 
turn of up to 40 pounds each of lunar material for Earth-based 
analysis. The SRC's are identical to the ones used on the 
Apollo 11 and 12 missions. They are machined from aluminum 
forgings and are designed to maintain an internal vacuum during 
the outbound and return flights. The SRC's will be filled with 
representative samples of lunar surface material, collected and 
separately bagged by the astronauts on their traverse and docu- 
mented by verbal descriptions and photography. Subsurface 
samples will be obtained by using drive tubes 16 inches long 
and one inch in diameter. A few grams of material will be 
preserved under lunar vacuum conditions in a special environ- 
mental sample container. 



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This container will be opened for analysis under vacuum 
conditions equivalent to that at the lunar surface. Special 
containers are provided for a magnetic sample and a gas analysis 
sample, 

SNAP-27 

SNAP-27 is one of a series of radioisotope thermoelectric 
generators, or atomic batteries, developed by the U.S. Atomic 
Energy Commission under its SNAP program. The SNAP (Systems 
for Nuclear Auxiliary Power) Program is directed at development 
of generators and reactors for use in space, on land, and in 
the sea. 

SNAP-27 was first used in the Apollo 12 mission to provide 
electricity for the first Apollo Lunar Surface Experiments 
Package (ALSEP). A duplicate of the Apollo 12 SNAP-27 will 
power the Apollo 13 ALSEP. 

The basic SNAP-27 unit is designed to produce at least 63 
electrical watts of power. It is a cylindrical generator 
fueled with the radioisotope plutonium 238, It is about 18 
inches high and 16 inches in diameter, including the heat radia- 
ting fins. The generator, making maximum use of the lightweight 
material beryllium, weighs about 28 pounds unfueled. 

The fuel capsule, made of a superalloy material, is 16.5 
inches long and 2.5 inches in diameter. It weighs about 15. 5 
pounds, of which 8.36 pounds represent fuel. The plutonium 
238 fuel is fully oxidized and is chemically and biologically 
inert . 

The rugged fuel capsule is contained within a graphite 
fuel cask from launch through lunar landing. The cask is de- 
signed to provide reentry heating protection and added contain- 
ment for the fuel capsule in the unlikely event of an aborted 
mission. The cylindrical cask with hemispherical ends includes 
a primary graphite heat shield, a secondary beryllium thermal 
shield, and a fuel capsule support structure made of titanium 
and Inconel materials. The cask is 23 inches long and eight 
inches in diameter and weighs about 24.5 pounds. With the fuel 
capsule installed, it weighs about HO pounds. It is mounted on 
the lunar module descent stage by a titanium support structure. 

Once the lunar module is on the Moon, the lunar module 
pilot will remove the fuel capsule from the cask and insert it 
into the SNAP-27 generator which will have been placed on the 
lunar surface near the module . 



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The spontaneous radioactive decay of the plutonium 238 
within the fuel capsule generates heat in the generator. 
An assembly of 442 lead telluride thermoelectric elements con- 
verts this heat — 1480 thermal watts — directly into elec- 
trical energy — at least 63 watts. There are no moving parts. 

Plutonium 238 is an excellent isotope for use in space 
nuclear generators. At the end of almost 90 years, plutonium 
238 will still supply half of its original heat. In the decay 
process, plutonium 238 emits mainly the nuclei of helium (alpha 
radiation), a very mild type of radiation with a short emission 
range . 

Before the use of the SNAP-27 system in the Apollo program 
was authorized, a thorough review was conducted to assure the 
health and safety of personnel involved in the mission and the 
general public. Extensive safety analyses and tests were con- 
ducted which demonstrated that the fuel would be safely con- 
tained under almost all credible accident conditions. 

Contractors for SNAP-27 

General Electric Co., Missile and Space Division, Phila- 
delphia, Pa., designed, developed, and fabricated the SNAP-27 
generator for the ALSEP. 

The 3M Co., St. Paul, Minn., fabricated the thermoelectric 
elements and assembled the SNAP-2 7 generator. 

Solar Division of International Harvester, San Diego, 
Calif., fabricated the generator's beryllium structure. 

Hitco, Gardena, Calif., fabricated the graphite structure 
for the SNAP-27 Graphite LM Fuel Cask. 

Sandia Corp., a subsidiary of Western Electric, operator 
of AEC T s Sandia Laboratory, Albuquerque s N.M., provided tech- 
nical direction for the SNAP-27 program. 

Savannah River Laboratory, Aiken, S.C., operated by the 
DuPont Co. for the AEC, prepared the raw plutonium fuel. 

Mound Laboratory, Miamisburg, Ohio, operated by Monsanto 
Research Corp., for the AEC, fabricated the raw fuel into the 
final fuel form and encapsulated the fuel. 



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PHOTOGRAFHIC EQUIPMENT 



Still and motion pictures will be made of most 
spacecraft maneuvers and crew lunar surface activities, 
and mapping photos from orbital altitude to aid in plan- 
ning future landing missions. During lunar surface activities, 
emphasis will be on photographic documentation of lunar 
surface features and lunar material sample collection. 

Camera equipment stowed in the Apollo 13 command 
module consists of two 70mm Hasselblad electric cameras, 
a l6mm motion picture camera, and the Hycon lunar topographic 
camera (LTC). 

The LTC, to be flown on Apollos 13, l 1 * and 15, is 
stowed beneath the commander's couch. In use, the camera 
mounts in -the crew access hatch window. 

The LTC with l8-inch focal length f/4.0 lens provides 
resolution of objects as small as 15-25 feet from a 60-nm 
altitude and as «mall as 3 to 5 feet from the 8-nm,pericynthion ( 
Film format Is 4.5-inch square frames on 100 foot long rolls s 
with a frame rate variable from k to 75 frames a minute. 
Shutter speeds are 1/50, 1/100, and 1/200 second. Spacecraft 
forward motion during exposures is compensated for by a servo- 
controlled rocking mount. The film is held flat in the focal 
plane by a vacuum platen connected to the auxiliary dump valve. 

The camera weighs 65 pounds without film, is 28 inches 
long, 10.5 inches wide, and 12.25 inches high. It is a mod- 
ification of an aerial reconnaissance camera. 

Future lunar landing sites and targets of scientific 
interest will be photographed with the lunar topographic 
camera in overlapping sequence of single frame modes. A 
candidate landing site northwest of the crater Censorinus 
will be photographed from the 8-mile pericynthion during the 
period between descent orbit insertion and CSM/LM separation. 
Additional topographic photos of the Censorinus site and 
sites near Davy Rille and Descartes will be made later in the 
mission from the 60-nm circular orbit. The camera again will 
be unstowed and mounted for 20 minutes of photography of the 
lunar disc at 5 minute Intervals starting at 2 hours after 
transearth injection. 



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Cameras stowed in the lunar module aro two 70mm 
Hasselblad data cameras fitted with 60mm Zeiss Metric lenses, 
a 16mm motion picture camera fitted with a 10mm lens, and a 
Kodak closeup stereo camera for high resolution photos on the 
lunar surface. The LM Hasselblads have crew chest mounts that 
leave both hands free. 

One of the command module Hasselblad electric cameras 
is normally fitted with an 80mm f/2.8 Zeiss Planar lens, but 
bayonet mount 250mm and 500irjn lenses may be substituted for 
special tasks. 

The second Hasselblad camera is fitted with an 80mm 
lens and a Reseau plate which allows greater dimensional 
control on photographs of the lunar surface. The 500mm lens 
will be used only as a backup to the lunar topographic camera. 

The 80mm lens has a focussing range from 3 feet to 
infinity and has a field of view of 38 degrees vertica.i and 
horizontal on the square-format film frame. Accessories for 
the command module Hasselblads include a spotmeter, inter- 

valometer, remote control cable, and film magazines . Hassel- 
blad shutter speeds range from time exposure and one second to 
one 1/500 second. 

The Maurer 16mm motion picture camera, in the command 
module has lenses of 5, 18, and 75mm available. The camera 
weighs 2.8 pounds with a 130-foot film magazine attached. 
Accessories include a right-angle mirror, a power cable, and 
a sestant adapter which allows the camera to use the navigation 
sextant optical system. The LM motion picture camera will be 
mounted in the right-hand window to record descent and landing 
and the two EVA periods and later will be taken to the surface. 

The 35 mm stereo closeup camera stowed in the 1M .MESA 
shoots 24mm square stereo pairs with an image scale of one- 
half actual size. The camera is fixed focus and is equipped 
with a stand-off hood to position the camera at the proper focus 
distance. A long handle permits an EVA crewman to position the 
camera without stooping for surface object photography. Detail 
as small as 40 microns can be recorded. The camera allows 
photography of significant surface structure which would remain 
intact only in the lunar environment, such as fine powdery 
deposits, cracks or holes, and adhesion of particles. A bat- 
tery-powered electronic flash provides illumination „ and film 
capacity is a minimum of 100 stereo pairs. 



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LUNAR DESCRIPTION 



Terrain - Mountainous and crater-pitted, the mountains 
rising as high as 29 thousand feet and the craters ranging 
from a few inches to 180 miles in diameter. The craters 
are thought to be formed primarily by the impact of 
meteorites. The surface is covered with a layer of fine- 
grained material resembling silt or sand, as well as small 
rocks and boulders. 

Environment - No air, no wind, and no moisture. The 
temperature ranges from 243 degrees P. in the two-week lunar 
day to 279 degrees below zero in the two-week lunar night. 
Gravity is one-sixth that of Earth. Micrometeoroids pelt 
the Moon since there is no atmosphere to burn them up. 
Radiation might present a problem during periods of unusual 
solar activity. 

Far Side - The far or hidden side of the Moon no longer 
is a complete mystery. It was first photographed by a 
Russian craft and since then has been photographed many 
times, .particularly from NASA's Lunar Orbiter and Apollo 
spacecraft. 

Origin - There is still no agreement among scientists 
on the origin of the Moon. The three theories: (1) the 
Moon once was part of Earth and split off into its own 
orbit, (2) it evolved as a separate body at the same time as 
Earth, and (3) it formed elsewhere in space and wandered 
until it was captured by Earth's gravitational field. 



Diameter 
Circumference 
Distance from Earth 

Surface temperature 

Surface gravity 

Mass 

Volume 

Lunar day and night 

Mean velocity in orbit 

Escape velocity 

Month (period of 
rotation around Earth) 



Physical Facts 

2,160 miles (about 1/4 that of Earth) 

6,790 miles (about 1/4 that of Earth) 

238,857 miles (mean; 221,463 minimum 
to 252,710 maximum) 

+243°F (Sun at zenith) -279°F (night) 

1/6 that of Earth 

l/100th that of Earth 

l/50th that of Earth 

14 Earth days each 

2,287 miles-per-hour 

1.48 miles-per-second 

27 days, 7 hours, 43 minutes 



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Landing Site 

The landing site selected for Apollo 13 Is located 
at 3° 40* 7"S, 17° 27 f 3"W, about 30 miles north of the 
Fra Mauro crater. The site is in a hilly, upland region. 
This will be the first Apollo landing to other than a lunar 
mare, the flat dark areas of the Moon once thought to be lunar 
seas. This hilly region has been designated as the Fra Mauro 
formation, a widespread geological unit covering large portions 
of the lunar surface around Mare Irabrium (Sea of Rains). The 
Fra Mauro formation is interpreted by lunar geologists to be 
an eject a blanket of material thrown out by the event which 
created the circular Mare Imbrium basin. 

The interpretation of the Fra Mauro formation as ejecta 
from Mare Imbrium gives rise to the expectation that surface 
material originated from deep within the Moon, perhaps from a 
hundred miles below the Moon's surface. If the interpretation 
proves correct, it will also be possible to date the Mare Imbrium 
event, believed to be a major impact, perhaps the in- fall of a 
smaller Moon, which was swept up in the primordial, accretionary 
evolution of the Moon. Based on this theory, rocks from the 
Fra Mauro formation should predate the rocks returned from 
either Apollo 11 (4.6 billion years) or Apollo 12 (3-5 billion 
years) and be close to the original age of the Moon, 



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AFOLLO 13 FLAGS, LUNAR MODULE PLAQUE 

The United States flag to be erected on the lunar 
surface measures 30 by 48 inches and will be deployed on 
a two-piece aluminum tube eight feet long. The folding 
horizontal bar which keeps the flag standing out from the 
staff on the airless Moon has been improved over the mech- 
anisms used on Apollo 11 and 12 . 

The flag, made of nylon, will be stowed in the lunar 
module descent stage modularized equipment stowage assembly 
(MESA) instead of in a thermal-protective tube on the LM front 
leg, as in Apollo 11 and 12. 

Also carried on the mission and returned to Earth will 
be 25 United States and 50 Individual state flags, each 4 by 
6 inches. 

A 7 by 9 inch stainless steel plaque, similar to those 
flown on Apollos 11 and 12, will be fixed to the LM front 
leg. The plaque has on it the words "Apollo 13" with "Aquarius" 
beneath, the date, and the signatures of the three crewmen. 



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SATURN" V LAUNCH VEHICLE 

The Saturn V launch vehicle (SA-508) assigned to the 
Apollo 13 mission was developed at the Marshall Space Flight 
Center, Huntsville, Ala. The vehicle is almost identical 
to those used in the missions of Apollo 8 through 12. 

First Stage 

The first stage (S-IC) of the Saturn V is built by the 
Boeing Company at NASA's Michoud Assembly Facility, New 
Orleans, La. The staged five F-l engines develop a total 
of about 7-6 million pounds of thrust at launch. Major com- 
ponents of the stage are the forward skirt, oxidizer tank, 
intertank structure, fuel tank, and thrust structure. Pro- 
pellant to the five engines normally flows at a rate of 
29,364.5 pounds (3s 400 gallons) each second. One engine is 
rigidly mounted on the stage's centerline; the other four 
engines are mounted on a ring at 90° angles around the center 
engine. These four outer engines are gimbaled to control the 
vehicle's attitude during flight. 

Second Stage 

The second stage (S-II) is built by the Space Division of 
the North American Rockwell Corporation at Seal Beach, Calif. 
Five J-2 engines develop a total of about 1.16 million pounds 
of thrust during flight. Major structural components are the 
forward skirt, liquid hydrogen and liquid oxygen tanks 
(separated by an insulated common bulkhead), a thrust structure, 
and an interstage section that connects the first and second 
stages. The five engines are mounted and used in the same way 
as the first stage's F-l engines: four outer engines can be 
gimbaled; the center one is rigid. 

Third Stage 

The third stage (S-IVB) is built by the McDonnell Douglas 
Astronautics Company at Huntington Beach, Calif. Major com- 
ponents are the aft interstage and skirt, thrust structure, 
two propellant tanks with a common bulkhead, a forward skirt, 
and a single J-2 engine. The gimbaled engine has a maximum 
thrust of 230,000 pounds, and can be shut off and restarted. 



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FIRST STAGE (S-IC) 
Diameter 33 feet 


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Fropellants -- Liquid oxygen (3,306, 1*9^ lbs.; 


3^0,3^3 gals.) RP-1 (kerosene) 


fl,i35,6U7 lbs.; 215,330 gala.) 


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Thrust 7,58U,593 lbs. at liftoff 






SECOND STAGE (S-II) 






THIRD STAGE 


Diameter 33 feet 






■ 1 (S-IVB) 


Height 81. 5 feet 






Weight 1,073,9^ lbs. fueled 








78,050 lbs. dry 


£ .111 




Engines Five J-2 




Propellants -- Liquid oxygen (836,120 lbs.; 


8 


SECOND STAGE 


88,215 gals.) liquid hydrogen 




(S-ll) 


(159,77^ lbs. ; £72,3^0 gals.) 


Thrust 92U,207 to 1,161,315 lbs. 


3: I— L. 

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Interstages—- 11,W>5 


THIRD STAGE (S-IVB) 


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Diameter 21.7 feet 


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Height 58.3 feet 




Weight 259,896 lbs. fueled 


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Propellants -- Liquid oxygen (191,532 Xbs.; 


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(^3,500 lbs., 6U,1U5 gals.) 








Thrust 199,790 lbs. 








Interstage 8,100 lbs. 




INSTRUMENT UNIT 












Height 3 feet 


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Weight k,k&2 lbs. 


. . 



NOTE: Weights and measures given above are for the nominal vehicle configura- 
tion for Apollo 12. The figures may vary slightly due to changes before launch 
to meet changing conditions. Weights of dry stages and propellants do not equal 
total weight because frost and miscellaneous smaller items are not included in 

chart. 

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Instrument Unit 



The instrument unit (IU), built by the International 
Business Machines Corp., at Huntsville, Ala., contains 
navigation, guidance and control equipment to steer the 
launch vehicle into its Earth orbit and into translunar 
trajectory. The six major systems are structural, thermal 
control, guidance and control, measuring and telemetry, 
radio frequency, and electric. 

The instrument unit provides a path-adaptive guidance 
scheme wherein a programmed trajectory is used during first 
stage boost with guidance beginning during second stage burn. 
This scheme prevents movements that could cause the vehicle 
to break up while attempting to compensate for winds or jet 
streams in the atmosphere. 

■ The instrument unit's inertial platform (heart of the 
navigation, guidance and control system) provides space-fixed 
reference coordinates and measures acceleration along three 
mutually perpendicular axes nf a coordinate system. If the 
platform fails during boost, systems in the Apollo spacecraft 
are programmed to provide guidance for the launch vehicle. 
After second stage ignition, the spacecraft commander could 
manually steer the vehicle in the event of loss of the launch 
vehicle inertial platform. 

Propulsion 

The Saturn V has 37 propulsive units, with thrust ratings 
ranging from 70 pounds to more than 1.5 million pounds. The 
large main engines burn liquid propellants; the smaller units 
use solid or hypergolic propellants. 

The five F-l engines on the first stage burn a combination 
of RP-1 (kerosene) as fuel and liquid oxygen as oxidizer. Each 
engine develops approximately 1,516,918 pounds of thrust at 
liftoff, building to about 1,799,022 pounds before cutoff. 
The five-engine cluster gives the first stage a thrust range 
of from 7,584,593 pounds at liftoff to 8,995,108 pounds just 
before center engine cutoff. The F-l engine weighs almost 
10 tons, is more than 18 feet long and has a nozzle exit 
diameter of nearly 14 feet. The engine consumes almost three 
tons of propellant every second. 

The first stage also has eight solid-fuel retrorockets 
that fire to separate the first and second stages. Each retro- 
rocket produces a thrust of 87,900 pounds for 0.6 seconds. 



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The second and third stages are powered by J-2 engines 
that burn liquid hydrogen (fuel) and liquid oxygen (oxidizer). 
J-2 engine thrust varies from 184,841 to 232,263 pounds 
during flight. The 3»500-pound J-2 engine is considered 
more efficient than the F-l engine because the J-2 burns high- 
energy liquid hydrogen. F-l and J-2 engines are built by 
the Rocketdyne Division of the North American Rockwell Corp, 

The second stage also has four 21 ,000-pound-thrust solid 
fuel ullage rockets that settle liquid propellant in the 
bottom of the main tanks and help attain a "clean" separation 
from the first stage. Four retrorockets, located in the S-IVB* s 
aft interstage (which never separates from the S-II)^ separate 
the S-II from the S-IVB. There are two jettisonable ullage 
rockets for propellant settling before engine ignition. Eight 
smaller engines in the two auxiliary propulsion system modules 
on the S-IVB stage provide three-axis attitude control. 



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COMMANP AND SERVICE MODULE STRUCTURE, SYSTEMS 

The Apollo spacecraft for the Apollo 13 mission Is comprised 
of Command Module 109, Service Module 109, Lunar Module 7> a 
spacecraft-lunar module adapter (SLA) and a launch escape system. 
The SLA houses the lunar module and serves as a mating structure 
between the Saturn V instrument unit and the SM. 

Launch Escape System (LES) — W° u J*- d propel command module to 
safety in an aborted launch. It has three solid-propellant 
rocket motors: a 147,000 pound- thrust launch escape system 
motor, a 2, 400 -pound-thrust pitch control motor, and a 31,500 
pound- thrust tower jettison motor. Two canard vanes deploy 
to turn the command module aerodynamically to an attitude with 
the heat-shield forward. The system is 33 feet tall and 4 feet, 
in diameter at the base, and weighs 8,945 pounds. 

Command Module (CM) Structure — The command module is a 
pressure vessel encased in heat shields, cone-shaped, weighing 
12,365 pounds at launch. 

The command module consists of a forward compartment which 
contains two reaction control engines and components of the Earth 
landing system; the crew compartment or inner pressure vessel 
containing crew accomodations, controls and displays, and many 
of the spacecraft systems; and the aft compartment housing ten 
reaction control engines, propellant tankage, helium tanks, water 
tanks, and the CSM umbilical cable. The crew compartment contains 
210 cubic feet of habitable volume. 

Heat-shields around the three compartments are made of 
brazed stainless steel honeycomb with an outer layer of phenolic 
epoxy resin as an ablative material. 

CSM and LM are equipped with the probe-and-drogue 
docking hardware. The probe assembly is a powered folding 
coupling and impact attentuating device mounted in the CM 
tunnel that mates with a conical drogue mounted in the LM 
docking tunnel. After the 12 automatic docking latches are 
checked following a docking maneuver, both the probe and 
drogue are removed to allow crew transfer between the CSM and 
LM. 

Service Module (SM) Structure — At launch, the service module 
for the Apollo 13 mission will weigh 51,105 pounds. Aluminum 
honeycomb panels one inch thick form the outer skin, and 
milled aluminum radial beams separate the interior into six 
sections around a central cylinder containing two helium spheres, 
four sections containing service propulsion system fuel-oxidizer 
tankage, another containing fuel cells, cryogenic oxygen and 
hydrogen, and one sector essentially empty. 



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Spacecraft-LM Adapter (SLA) Structure — The spacecraft LM 
adapter is a truncated cone 2 8 feet long tapering from 260 inches 
diameter at the base to 154 inches at the forward end at the 
service module mating line. The SLA weighs 4,000 pounds and 
houses the LM during launch and Earth orbital flight. 

CSM Systems 

Guidance, Navigation and Control System (GNCS) — Measures 
and controls spacecraft position, attitude, and velocity, 
calculates trajectory, controls spacecraft propulsion system 
thrust vector, and displays abort data. The guidance system 
consists of three subsystems: Inertial, made up of an inertial 
measurement unit and associated power and data components; 
computer which processes information to or from other components; 
and optics consisting of scanning telescope and sextant for 
celestial and/or landmark sighting for spacecraft navigation. 
VHF ranging device serves as a backup to the LM rendezvous radar. 

Stabilization and Control Systems (SCS) — Controls space- 
craft rotation, translation, and thrust vector and provides 
displays for crew-initiated maneuvers; backs up the guidance system 
for control functions. It has three subsystems; attitude 
reference, attitude control, and thrust vector control. 

Service Propulsion System (SPS) — Provides thrust for large 
spacecraft velocity changes through a gimbal-mounted 20, 500-pound- 
thrust hypergolic engine, using a nitrogen tetroxide oxidizer and 
a 50-50 mixture of unsymmetrical dimethyl hydrazine and hydrazine 
fuel. This system is in the service module. The system responds 
to automatic firing commands from the guidance and navigation 
system or to manual commands from the crew. The engine thrust 
level is not throttleable . The stabilisation and control 
system gimbals the engine to direct the thrust vector through the 
spacecraft center of gravity. 

Telecommunications System — Provides voice, television,' 
telemetry, and command data and tracking and ranging between 
the spacecraft and Earth, between the command module and the 
lunar module and between the spacecraft and astronauts during 
EVA. It also provides intercommunications between astronauts. 

The high-gain steerable S-Band antenna consists of four, 31- 
inch-diameter parabolic dishes mounted on a folding boom at the 
aft end of the service module. Signals from the ground stations 
can be tracked either automatically or manually with the antenna's 
gimballing system. Normal S-Band voice and uplink/downlink 
communications will be handled by the omni and high-gain antennas. 



■82- 



Sequential System — Interfaces with other spacecraft systems 
and subsystems to initiate time critical functions during launch, 
docking maneuvers, sub-orbital aborts, and entry portions of a 
mission. The system also controls routine spacecraft sequencing 
such as service module separation and deployment of the Esrth 
landing system. 

Emergency Detection System (EDS) — Detects and displays to 
the crew launch vehicle emergency conditions, such as excessive 
pitch or roll rates or two engines out , and automatically or 
manually shuts down the booster and activates the launch escape 
system; functions until the spacecraft is in orbit. 

Earth Landing System (ELS) — Includes the drogue and main 
parachute system as well as post-landing recovery aids. In a 
normal entry descent, the command module forward heat shield 
is jettisoned at 24,000 feet, permitting mortar deployment of 
two reefed 16 . 5-foot diameter drogue parachutes for orienting 
and decelerating the spacecraft. After disreef and drogue release, 
three mortar deployed pilot chutes pull out the three main 83.3- 
foot diameter parachutes with two-stage reefing to provide gradual 
inflation in three steps. Two main parachutes out of three can 
provide a safe landing. 

reaction Control System (RCS) — The SM RCS has four identical 
RCS "quads" mounted around the SM 90 degrees apart. Each quad 
has four 100 pound-thrust engines, two fuel and two oxidizer tanks 
and a helium pressurization sphere. Attitude control and small 
velocity maneuvers are made with the SM RCS. 

The CM RCS consists of two independent six-engine subsystems 
of six 9 3 pound-thrust engines each used for spacecraft attitude 
control during entry. Propellants for both CM and SM RCS are 
monomethyl hydrazine fuel and nitrogen tetroxide oxidizer with 
helium pre ssurizati on . These propellants burn spontaneously 
when combined (without an igniter) . 

Electrical Power System (EPS) — Provides electrical energy 
sources, power generation and control, power conversion i 
conditioning, and distribution to the spacecraft. The 
primary source of electrical power is the fuel cells mounted in 
the SM. The fuel cell also furnishes drinking water to the 
astronauts as a by-product. 



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Three silver-zinc oxide storage batteries supply power to 
the CM during entry and after landing, provide power for sequence 
controllers, and supplement the fuel cells during periods of 
peak power demand. A battery charger assures a full charge prior 

to entry . 

Two other silver-zinc oxide batteries supply power for 
explosive devices for CM/SM separation, parachute deployment 
and separation, third-stage separation, launch escape 
tower separation, and other pyrotechnic uses. 

Environmental Control System (ECS) — Controls spacecraft 
atmosphere, pressure, and temperature and manages water. In 
addition to regulating cabin and suit gas pressure, temperature 
and humidity, the system removes carbon dioxide, odors and 
particles and ventilates the cabin after landing. It collects 
and stores fuel cell potable water for crew use, supplies water 
to the glycol evaporators for cooling, and dumps surplus water 
overboard through the waste H 2 dump nozzle. Proper operating 
temperature of electronics ana electrical equipment is maintained 
by this system through the use of the cabin heat exchangers, the 
space radiators, and the glycol evaporators. 

Recovery Aids — Recovery aids include the uprighting 
system, swimmer interphone connections, sea dye marker, flash- 
ing beacon, VHF recovery beacon, and VHP transceiver. The up- 
righting system consists of three compressor-inflated bags to 
upright the spacecraft if it should land in the water apex 
down (stable II position). 

Caution and Warning System — Monitors spacecraft systems 
for out-of-tolerance conditions and alerts crew by visual and 
audible alarms. 

Controls and Displays — Provide status readouts and 
control functions of spacecraft systems in the command and 
service modules. All controls are designed to be operated by 
crewmen in pressurized suits. Displays are grouped by system 
and located according to the frequency of use and crew responsibility 



-84- 

LUNAR MODULE STRUCTURES, WEIGHT 

The lunar module is a two-stage vehicle designed for 
space operations near and on the Moon. The lunar module stands 
22 feet 11 inches high and is 31 feet wide (diagonally across 
landing gear). The ascent and descent stages of the LM operate as 
a unit until staging, when the ascent stage functions as a 
single spacecraft for rendezvous and docking with the CM. 

Ascent Stage 

Three main sections make up the ascent stage: the crew 
compartment, midsection, and aft equipment bay. Only the crew 
compartment and midsection are pressurized (4.8 psig) . The 
cabin volume is 2 35 cubic feet (6.7 cubic meters). The stage 
measures 12 feet 4 inches high by 14 feet 1 inch in diameter. 
The ascent stage has six substructural areas: crew compartment, 
midsection, aft equipment bay, thrust chamber assembly cluster 
supports, antenna supports, and thermal and micrometeoroid shield. 

The cylindrical crew compartment is 92 inches (2.35 m) in 
diameter and 42 inches (1.07 m) deep. Two flight stations are 
equipped with control and display panels, armrests, body restraints, 
landing aids, two front windows, an overhead docking window, and 
an alignment optical telescope in the center between the two 
flight stations. The habitable volume is 160 cubic feet. 

A tunnel ring atop the ascent stage meshes with the 
command module docking latch assemblies. During docking, the 
CM docking ring and latches are aligned by the LM drogue and 
the CSM probe. 

The docking tunnel extends downward into the midsection 
16 inches (40 cm). The tunnel is 32 inches (81 cm) in 
diameter and is used for crew transfer between the CSM and LM. 
The upper hatch on the inboard end of the docking tunnel opens 
inward and cannot be opened without equalizing pressure on both 
hatch surfaces. 

A thermal and micrometeoroid shield of multiple layers 
of Mylar and a single thickness of thin aluminum skin encases 
the entire ascent stage structure. 



-85- 



DOCKING WINDOW 



S-BAND 
STEERABLE 
ANTENNA 

RENDEZVOUS 
RADAR ANTENNA 

S-BAND IN-FUGHT 
ANTENNA (2) 



WINDOWS p) 



TRACKING LIGHT 



DOCKING 

DROGUE 

ASSEMBLY 



DOCKING 
TARGET 



EVA ANTENNA 



AFT 

EQUIPMENT 

BAY 

RCS THRUST 
CHAMBER 
ASSEMBLY 
CLUSTER (4) 




DOCKING 
LIGHT (4) 



RTG CASK 

LANDING 
PAD 



LUNAR SURFACE SENSING PROBE (3) 



APOLLO LUNAR MODULE 



-mo re - 



-86- 

Descent Stage 

The descent stage center compartment houses the descent 
engine, and descent propellant tanks are housed In the four 
square bays around the engine. Quadrant II (Seq bay) contains 
ALSEP, and Radioisotope Thermoelectric Generator (RTG) externally. 
Quadrant IV contains the MESA. The descent stage measures 10 
feet 7 inches high by 14 feet 1 Inch In diameter and is encased 
in the Mylar and aluminum alloy thermal and micrometeoroid shield. 

The LM egress platform, or "porch", is mounted on the forward 
outrigger just below the forward hatch. A ladder extends down the 
forward landing gear strut from the porch for crew lunar surface 
operations . 

The landing gear struts are explosively extended and provide 
lunar surface landing impact attenuation. The main struts are 
filled with crushable aluminum honeycomb for absorbing 
compression loads. Footpads 37 inches (0.95 m) In diameter at 
the end or each landing gear provide vehicle support on the 
lunar surface . 

Each pad (except forward pad) is fitted with a 68 Inch 
long lunar surface sensing probe which signals the crew to shut 
down the descent engine upon contact with the lunar surface. 

LM-7 flown on the Apollo 13 mission has a launch weight of 
33,^76 pounds. The weight breakdown is as follows: 

Ascent stage, dry M,668 lbs. Includes water 

and oxygen; no 
Descent stage, dry 4,650 lbs. crew 

RCS propellants (loaded) 590 lba. 

DPS propellants (loaded) 18,339 lbs, 

APS propellants (loaded) 5^229 lba. 

33,^76 lbs. 

Lunar Module Systems 

Electrical Power System — The LM DC electrical system consists 
of six silver zinc primary batteries — four in the descent stage 
and two in the ascent stage. Twenty-eight-volt DC power is 
distributed to all LM systems. AC power (117v 400 Hz) Is supplied 
by two inverters. 



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Environmental Control System — Consists of the atmosphere 
revitalization section, oxygen supply and cabin pressure control 
section, water management, heat transport section, and outlets 
for oxygen and water reservicing of the portable life support 
system (PLSS). 

Components of the atmosphere revitalization section are the 
suit circuit assembly which cools and ventilates the pressure 
garments, reduces carbon dioxide levels, removes odors, noxious 
gases and excessive moisture; the cabin recirculation assembly 
which ventilates and controls cabin atmosphere temperatures; and 
the steam flex duct which vents to space steam from the suit 
circuit water evaporator. 

The oxygen supply and cabin pressure section supplies gaseous 
oxygen to the atmosphere revitalization section for maintaining 
suit and cabin pressure. The descent stage oxygen supply provides 
descent flight phase and lunar stay oxygen needs, and the ascent 
stage oxygen supply provides oxygen needs for the ascent and 
rendezvous flight phase. 

Water for drinking, cooling, fire fighting, food preparation, 
and refilling the PLSS cooling water servicing tank is supplied by 
the water management section. The water is contained in three 
nitrogen-pressurized bladder-type tanks, one of 367-pound capacity 
in the descent stage and two of 47.5-pound capacity in the ascent 
stage . 

The heat transport section has primary and secondary water- 
glycol solution coolant loops. The primary coolant loop circulates 
water-glycol for temperature control of cabin and suit circuit 
oxygen and for thermal control of batteries and electronic compon- 
ents mounted on cold" plates and rails. If the primary loop becomes 
inoperative, the secondary loop circulates coolant through the 
rails and cold plates only. Suit circuit cooling during secondary 
coolant loop operation is provided by the suit loop water boiler. 
Waste heat from both loops is vented overboard by water evaporation 
or sublimators. 

Communications System — Two S-band transmitter-receivers, 
two VHF transmitter-receivers, a signal processing assembly, 
and associated spacecraft antenna make up the LM communications 
system. The system transmits and receives voice and tracking and 
ranging data, and transmits telemetry data on about 270 measure- 
ments and TV signals to the ground. Voice communications 
between the LM and ground stations is by S-band, and between the 
LM and CSM voice is on VHF. 

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Although no real-time commands can be sent to the LM, 
the digital uplink processes guidance officer commands, such 
as state vector updates, transmitted from Mission Control 
Center to the LM guidance computer. 

The data storage electronics assembly (DSEA) is a 
four-channel voice recorder with timing signals, with a 
10-hour recording capacity, which will be brought back into 
the CSM for return to Earth. DSEA recordings cannot be 
"dumped" to ground stations . 

LM antennas are one 26-inch-diameter parabolic S-band 
steerable antenna, two S-band inflight antennas, two VHF in- 
flight antennas, EVA antenna, and an erectable S-band antenna 
(optional) for lunar surface. 

Guidance, Navigation, and Control System — Comprised 
of six sections: primary guidance and navigation section (PGNS), 
abort guidance section (AGS), radar section, control electron- 
ics section (CES), and orbit rate display Earth and lunar 
(ORDEAL). 

* The PGNS is an aided inertial guidance system updated 
by the alignment optical telescope, an inertial measurement 
unit, and the rendezvous and landing radars. The system pro- 
vides inertial reference data for computations, produces in- 
ertial alignment reference by feeding optical sighting data into 
the LM guidance computer, displays position and velocity data, 
computes LM-CSM rendezvous data from radar inputs, controls 
attitude and thrust to maintain desired LM trajectory, and 
controls descent engine throttling and gimbaling. 

The LM-7 primary guidance computer has the Luminary 1C 
Software program, which is an improved version over that in 

LM-6. 

* The AGS is an independent backup system for the PGNS, 
having its own inertial sensors and computer. 

* The radar section is made up of the rendezvous radar 
which provides CSM range and range rate, and line-of-sight 
angles for maneuver computation to the LM guidance computer; 
and the landing radar which provides altitude and velocity 
data to the LM guidance computer during lunar landing. The 
rendezvous radar has an operating range from 80 feet to 400 
nautical miles. The ranging tone transfer assembly, utilizing 
VHF electronics, is a passive responder to the CSM VHF ranging 
device and is a backup to the rendezvous radar. 



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* The CES controls LM attitude and translation about all 
axes. It also controls by PGNS command the automatic operation 
of the ascent and descent engine and the reaction control thrusters. 
Manual attitude controller and thrust-translation controller 
commands are also handled by the CES. 

*ORDEAL, displayed on the flight director attitude indicator, 
is the computed local vertical in the pitch axis during circular 
Earth or lunar orbits . 

Reaction Control System — The LM has four RCS engine clusters 
of four 100-pound (45.** kg) thrust engines each, which use helium- 
pressurized hypergolic propellants. The oxidizer Is nitrogen 
tetroxide, fuel is Aerozine 50 (50/50 blend of hydrazine and 
unsymmetrical dimethyl hydrazine). Interconnect valves permit the 
RCS system to draw from ascent engine propellant tanks. 

The RCS provides small stabilizing impulses during ascent and 
descent burns, controls LM attitude during maneuvers, and produces 
thrust for separation, and for ascent/descent engine tank ullage. 
The system may be operated in either the pulse or steady-state modes. 

Descent Propulsion System — Maximum rated thrust of the 
descent engine is 9>870 pounds (4,380.9 kg) and is throttleable 
between 1,050 pounds (476.7 kg) and 6,300 pounds (2,860.2 kg). 
The engine can be gimbaled six degrees in any direction in response 
to attitude commands and to compensate for center of gravity offsets 
Propellants are helium-pressurized Aerozine 50 and nitrogen 
tetroxide. 

Ascent Propulsion System — The 3,500-pound (1,589 kg) 
thrust ascent engine is not gimbaled and performs at full thrust. 
The engine remains dormant until after the ascent stage is separated 
from the descent stage. Propellants are the same as are burned 
by the RCS engines and the descent engine. 

Caution and Warning, Controls and Displays — These two systems 
have the same function aboard the lunar module as they do aboard the 
command module (See CSM systems section.) 

Tracking and Docking Lights -- A flashing tracking light (once 
per second, 20 milliseconds duration) on the front face of the 
lunar module is an aid for contingency CSM-active rendezvous LM 
rescue. Visibility ranges from 400 nautical miles through the CSM 
sextant to 130 miles with the naked eye. Five docking lights 
analagous to aircraft running lights are mounted on the LM for 
CSM-active rendezvous: two forward yellow lights, aft white light, 
port red light and starboard green light. All docking lights have 
about a 1,000-foot visibility. 



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APOLLO 13 CREW AND CREW EQUIPMENT 

Life Support Equipment - Space Suits 

Apollo 13 crewmen will wear two. versions of the Apollo 
space suit: an intravehicular pressure garment assembly 
worn by the command module pilot and the extravehicular 
pressure garment assembly worn by the commander and the lunar 
module pilot. Both versions are basically identical except 
that the extravehicular version has an integral thermal/ 
meteoroid garment over the basic suit. 

Prom the skin out, the basic pressure garment consists 
of a nomex comfort layer, a neoprene-coated nylon pressure 
bladder and a nylon restraint layer. The outer layers of the 
intravehicular suit are, from the inside out, nomex and two 
layers of Teflon-coated Beta cloth. The extravehicular inte- 
gral thermal /meteoroid cover consists of a liner of two layers 
of neoprene-coated nylon, seven layers of Beta/Kapton spacer 
laminate, and an outer layer of Teflon-coated Beta fabric. 

The extravehicular suit, together with a liquid cooling 
garment, portable life support system (PLSS), oxygen purge 
system, lunar extravehicular visor assembly and other components 
make up the extravehicular mobility unit (EMU). The EMU pro- 
vides an extravehicular crewman with life support for a four- 
hour mission outside the lunar module without replenishing 
expendables. EMU total weight is 183 pounds. The intra- 
vehicular suit weighs 35.6 pounds. 

Liquid cooling garment — A knitted nylon-spandex garment 
with a network of plastic tubing through which cooling water 
from the PLSS is circulated. It is worn next to the skin and 
replaces the constant wear-garment during EVA only. 

Portable life support system — A backpack supplying oxygen 
at 3*9 psi and cooling water to the liquid cooling garment. 
Return oxygen is cleansed of solid and gas contaminants by a 
lithium hydroxide canister. The PLSS includes communications 
and telemetry equipment, displays and controls, and a main 
power supply. The PLSS is covered by a thermal insulation 
Jacket. (Two stowed in LM) . 

Oxygen purge system — Mounted atop the PLSS, the oxygen 
purge system provides a contingency 45-minute supply of 
gaseous oxygen in two two-pound bottles pressurized to 5,880 
psia. The system may also be worn separately on the front of 
the pressure garment assembly torso. It serves as a mount for 
the VHF antenna for the PLSS. (Two stowed in LM) . 

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BACKPACK SUPPORT STRAPS 



OXYGEN PURGE SYSTEM 



SUNGLASSES 
POCKET 



PORTABLE LIFE 
SUPPORT SYSTEM 



OXYGEN 

PURGE SYSTEM 

UMBILICAL 



LM RESTRAINT RING 

INTEGRATED THERMAL 
METEOROID GARMENT 



URINE TRANSFER CONNECTOR, 
BIOMEDICAL INJECTION, 
DOSIMETER ACCESS FLAP AND 
DONNING LANYARD POCKET 



LUNAR EXTRAVEHICULAR VISOR 



Hl\ BACKPACK CONTROL BOX 



OXYGEN PURGE 
SYSTEM ACTUATOR 



PENLIGHT POCKET 
CONNECTOR COVER 
COMMUNICATION, 
VENTILATION, AND 
LIQUID COOLING 
UMBILICALS 




EXTRAVEHICULAR 
GLOVE 

UTILITY POCKET 



LUNAR OVERSHOE 



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EXTRAVEHICULAR MOBILITY UNIT 



-92- 



Lunar extravehicular visor assembly — A polycarbonate 
shell and two visors with thermal control and optical 
coatings on them. The EVA visor is attached over the 
pressure helmet to provide impact, micrometeoroid , thermal 
and ultraviolet-infrared light protection to the EVA crew- 
men. Since Apollo 12, a sunshade has been added to the outer 
portion of the LEWA in the middle portion of the helmet rim. 

Extravehicluar gloves — Built of an outer shell of 
Chromel-R fabric and thermal insulation to provide protection 
when handling extremely hot and cold objects. The finger 
tips are made of silicone rubber to provide more sensitivity. 

A one-piece constant-wear garment , similar to "long 
Johns," is worn as an undergarment for the space suit in intra- 
vehicular operations and for the inflight coveralls. The 
garment is porous-knit cotton with a waist-to-neck zipper for 
donning. Biomedical harness attach points are provided. 

During periods out of the space suits, crewmen wear two- 
piece Teflon fabric inflight coveralls for warmth and for 
pocket stowage of personal items. 

Communications carriers ("Snoppy Hats") with redundant 
microphones and earphones are worn with the pressure helmet; 
a lightweight headset is worn with the inflight coveralls. 

Another modification since Apollo 12 has been the addition 
of eight-ounce drinking water bags ("Gunga Dins") attached to 
the inside neck rings of the EVA suits. The crewmen can take 
a sip of water from the 6 X 8 inch bag through a 1/8-inch- 
diameter tube within reach of his mouth. The bags are filled 
from the lunar module potable water dispenser. 



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Apollo Lunar Hand Tools 

Special Environmental Container - The special environ- 
mental sample Is collected in a carefully selected area and 
sealed in a special container which will retain a high vacuum. 
The container is opened in the Lunar Receiving Laboratory where 
it will provide scientists the opportunity to study lunar 
material in its original environment. 

Extension handle - This tool is of aluminum alloy tubing 
with a malleable stainless steel cap designed to be used as an 
anvil surface. The handle is designed to be used as an extension 
for several other tools and to permit" their use without re- 
quiring the astronaut to kneel or bend down. The handle is 
approximately 24 inches long and 1 inch in diameter. The 
handle contains the female half of a quick disconnect fitting 
designed to resist compression, tension, torsion, or a combina- 
tion of these loads. 

Three core tubes - These tubes are designed to be driven 
or augered into loose gravel, sandy material, or into soft rock 
such as feather rock or pumice. They are about 15 inches in 
length and one inch in diameter and are made of aluminum 
tubing. Each tube is supplied with a removeable non-serrated 
cutting edge and a screw-on cap incorporating a metal-to-metal 
crush seal which replaces the cutting edge. The upper end of 
each tube is sealed and designed to be used with the extension 
handle or as an anvil. Incorporated into each tube is a spring 
device to retain loose materials in the tube. 

Scoops (large and small) - These tools are designed for 
use as a trowel and as a chisel. The scoop is fabricated 
primarily of aluminum with a hardened-steel cutting edge 
riveted on and a nine-inch handle. A malleable stainless steel 
anvil is on the end of the handle. The angle between the 
scoop pan and the handle allows a compromise for the dual use. 
The scoop is used either by itself or with the extension 
handle. The large scoop has a seive which permits particles 
smaller than 1/2 cm to pass through. 

Sampling hammer - This tool serves three functions, as a 
sampling hammer, as a pick or mattock, and as a hammer to 
drive the core tubes or scoop. The head has a small hammer face 
on one end, a broad horizontal blade on the other, and large 
hammering flats on the sides. The handle is 14 inches long and 
is made. of formed tubular aluminum. The hammer has on its 
lower end a quick-disconnect to allow attachment to the exten- 
sion handle for use as a hoe. The head weight has been in- 
creased to provide more Impact force. 



-96- 

BRUSH/ SCRIBER/ HAND LENS 





CORE TUBE 
AND CAP 




TONGS 




GEOLOGIC SAMPLING TOOLS 



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-97- 



Tongs - The tongs are designed to allow the astronaut to 
retrieve small samples from the lunar surface while in a 
standing position. The tines are of such angles, length, and 
number to allow samples of from 3/8 up to 2-1/2-inch diameter 
to be picked up. This tool is 24 inches in overall length. 

Brush/Scrlber/Hand Lens - A composite tool 

(1) Brush - To clean samples prior to selection 

(2) Scriber - To scratch samples for selection and to 
mark for identification 

(3) Hand lens - Magnifying glass to facilitate sample 
selection 

Spring Scale - To weigh two rock boxes and other bags 
containing lunar material samples, to maintain weight budget 
for return to Earth. 

Instrument staff - The staff hold the Hasselblad camera. 
The staff breaks down into sections. The upper section telescopes 
to allow generation of a vertical stereoscopic base of one foot 
for photography. Positive stops are provided at the extreme of 
travel. A shaped hand grip aids in aiming and carrying. The 
bottom section is available in several lengths to suit the staff 
to astronauts of varying sizes. The device is fabricated from 
tubular aluminum. 

Gnomon - This tool consists of a weighted staff suspended 
on a two-ring gimbal and supported by a tripod. The staff 
extends 12 inches above the gimbal and is painted with a gray 
scale. The gnomon is used as a photographic reference to 
indicate local vertical, sun angle, and scale. The gnomon has a 
required accuracy of vertical indication of 20 minutes of arc. 
Magnetic damping is incorporated to reduce oscillations. 

Color Chart - The color chart is painted with three primary 
colors and a gray scale. It is used as a calibration for lunar 
photography. The scale Is mounted on the tool carrier but may 
easily be removed and returned to Earth for reference. The color 
chart is 6 inches in size. 



-more- 



-98- 

Tool Carrier - The carrier is the stowage container for 
the tools during the lunar flight. After the landing the 
carrier serves as support for the astronaut when he kneels 
down, as a support for the sample bags and samples, and as 
a tripod base for the instrument staff. The carrier folds 
flat for stowage. For field use It opens into a triangular 
configuration. The carrier Is constructed of formed sheet 
metal and approximates a truss structure. Six-inch legs 
extend from the carrier to elevate the carrying handle suffi- 
ciently to be easily grasped by the astronaut. 

Field Sample Bags - Approximately 80 bags four Inches by 
five Inches are included in the Apollo lunar hand tools for 
the packaging of samples. These bags are fabricated from 
Teflon FEP. 

Collection Bag - This is a large bag (H X 8 inches) 
attached to the astronaut's side of the tool carrier. Field 
sample bags are stowed in this bag after they have been filled, 
It can also be used for general storage or to hold items 
temporarily. (Two in each SRC). 

Trenching Tool - A trenching tool with a pivoting scoop 
has been provided for digging the two-foot deep soil mechanics 
investigation trench. The two-piece handle Is five feet long. 
The scoop is eight Inches long and five inches wide and pivots 
from in-line with the handle to 90° — similar to the trenching 
tool carried on infantry backpacks. The trenching tool is 
stowed in the MESA rather than in the tool carrier. 

Lunar Surface Drill - The 29.4-pound Apollo Lunar Surface 
Drill (ALSD) Is stowed in the ALSEP subpackage No. 2 and will 
be used .for boring two ten-foot deep 1.25-inch diameter holes 
for ALSEP heat flow experiment probes, and one approximately 
eight-foot-deep, one -inch- diameter core sample. The silver- 
zinc battery-powered rotary percussive drill has a clutch to 
limit torque to 20 foot-pounds. A treadle assembly serves as 
a drilling platform and as a core stem lock during the drill 
string decoupling operation as the string is withdrawn from 
the lunar soil. Bore stems for the heat flow experiment holes 
are of boron/fiberglas , and the core sample core stems are 
titanium. Cutting bits are tungsten carbide. 



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-99- 

Apollo 13 Crew Menu 

More than 70 items comprise the food selection list of 
freeze-dried rehydratable, wet-pack and spoon-bowl foods. 
Balanced meals for five days have been packed in man/day 
overwraps. Items similar to those in the daily menus have 
been packed in a snack pantry. The snack pantry permits the 
crew to locate easily a food item in a smorgasbord mode with- 
out having to "rob" a regular meal somewhere down deep in a 
storage box. 

Water for drinking and rehydrating food is obtained from 
two sources in the command module — a dispenser for drinking 
water and a water spigot at the food preparation station sup- 
plying water at about 155 or 55° P. The potable water dis- 
penser squirts water continuously as long as the trigger is 
held down, and the food preparation spigot dispenses water 
in one-ounce increments . 

A continuous-feed hand water dispenser similar to the 
one in the command module is used aboard the lunar module 
for cold-water rehydration of food packets stowed aboard the 
LM. 

After water has been injected into a food bag, it is 
kneaded for about three minutes. The bag neck is then cut 
off and the food squeezed into the crewman's mouth. After 
a meal, germicide pills attached to the outside of the food 
bags are placed in the bags to prevent fermentation and gas 
formation. The bags are then rolled and stowed in waste 
disposal compartments. 

The day-by-day, meal-by-meal Apollo 13 Menu for Com- 
mander Lovell is on the following page as a typical five- 
day menu for each crewman. 



-more- 



-100- 



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-101- 



Personal Hygiene 

Crew personal hygiene equipment aboard Apollo 13 
includes body cleanliness items, the waste management system 
and one medical kit. 

Packaged with the food are a toothbrush and a two-ounce 
tube of toothpaste for each crewman. Each man-meal package 
contains a 3«5-by-4-inch wet-wipe cleansing towel. Addition- 
ally, three packages of 12-by-12-inch dry towels are stowed 
beneath the command module pilot's couch. Each package con- 
tains seven towels. Also stowed under the command module 
pilot's couch are seven tissue dispensers containing 53 three- 
ply tissues each. 

Solid body wastes are collected in plastic defecation 
bags which contain a germicide to prevent bacteria and gas 
formation. The bags are sealed after use and stowed in empty 
food containers for post-flight analysis. 

Urine collection devices are provided for use while 
wearing either the pressure suit or the inflight coveralls. 
The urine is dumped overboard through the spacecraft urine dump 
valve in the CM and stored in the LM. 

Medical Kit 

The 5X5X8-inch medical accessory kit is stowed in a com- 
partment on the spacecraft right side wall beside the lunar 
module pilot couch. The medical kit contains three motion 
sickness injectors, three pain suppression injectors, one two- 
ounce bottle first aid ointment, two one-ounce bottles eye 
drops, three nasal sprays, two compress bandages, 12 adhesive 
bandages, one oral thermometer, and four spare crew biomedical 
harnesses. Pills in the medical kit are 60 antibiotic, 12 
nausea, 12 stimulant, 18 pain killer, 60 decongestant, 24 
diarrhea, 72 aspirin and 21 sleeping. Additionally, a small 
medical kit containing four stimulant, eight diarrhea, two 
sleeping and four pain killer pills, 12 aspirin, one bottle 
eye drops, two compress bandages, 8 decongestant pills, one 
pain injector, one bo.ttle nasal spray is stowed in the lunar 
module flight data file compartment. 



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-102- 



Survival Gear 

The survival kit is stowed in two rucksacks in the 
right-hand forward equipment bay above the lunar module 
pilot. 

Contents of rucksack No. 1 are: two combination sur- 
vival lights, one desalter kit, three pair sunglasses, one 
radio beacon, one spare radio beacon battery and spacecraft 
connector cable, one knife in sheath, three water containers, 
and two containers of Sun lotion, two utility knives, three 
survival blankets and one utility netting. 

Rucksack No. 2: one three-man life raft with COg 
inf later, one sea anchor, two sea dye markers, three sun- 
bonnets, one mooring lanyard, three manlines and two attach 
brackets. 

The survival kit is designed to provide a 48-hour post- 
fending (water or land) survival capability for three crewmen 
between 40° North and South latitudes. 

Biomedical Inflight Monitoring 

The Apollo 13 crew biomedical telemetry data received 
by the Manned Space Flight Network will be relayed for in- 
stantaneous display at Mission Control Center where heart 
rate and breathing rate data will be displayed on the flight 
surgeon's console. Heart rate and respiration rate average, 
range and deviation are computed and displayed on digital TV 
screens . 

In addition, the instantaneous heart rate, real-time and 
delayed EKG and respiration are recorded on strip charts for 
each man. 

Biomedical telemetry will be simultaneous from all crew- 
men while in the CSM, but selectable by -a manual onboard 
switch in the LM. 

Biomedical data observed by the flight surgeon and his 
team in the Life Support Systems Staff Support Room will be 
correlated with spacecraft and space suit environmental data 
displays. 

Blood pressures are no longer telemetered as they were in 
the Mercury and Gemini programs. Oral temperatures, however, 
can be measured onboard for diagnostic purposes and voiced 
down by the crew in case of inflight illness. 



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-103* 



Energy expended by the crewmen during EVA will be 
determined indirectly using a metabolic computation pro- 
gram based on three separate measurements: 

1) Heart rate portion — Heart rate will be determined 
from telemetered EKG and converted to oxygen con- 
sumption (litre/min) and heat production (BTU/hour) 
based on pre-f light calibration curves. These curves 
are determined from exercise response tests utilizing 
a bicycle ergometer. 

2) Oxygen usage portion — Oxygen usage will be determined 
from the telemetered measurement of PLSS oxygen supply 
pressure. Suit leak determined pre-f light is taken 
into account. Heat production will be calculated from 
oxygen usage. 

3) Liquid cooled garment temperature portion — The 
amount of heat taken up by the liquid cooled garment 
will be determined from telemetered measurements of 
the LCG water temperature inlet and change in/out. 
This measurement (the amount of heat taken up by the 
water) plus an allowance made for sensible and latent 
heat loss, radiant heat load, and possible heat storage 
will provide an indication of heat production by the 
crewman. 



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-104- 
Training 

: The crewmen of Apollo 13 have spent more than five hours 
of formal crew training for each hour of the lunar- launching 
mission's ten-day duration. More than 1,000 hours of 
training were in Apollo 13 crew training syllabus over and 
above the normal preparations for the mission — technical 
briefings and reviews, pilot meetings and study. 

The Apollo 13 crewmen also took part in prelaunch 
testing at Kennedy Space Center, such as altitude chamber 
tests and the countdown demonstration tests (CDDT) which 
provided the crew with thorough operational knowledge of 
the complex vehicle. 



Highlights of specialized Apollo 13 crew training topics 
are; 

* Detailed series of briefings on spacecraft systems, 
operation and modifications. 

* Saturn launch vehicle briefings on countdown, range 
safety, flight dynamics, failure modes and abort conditions. 
The launch vehicle briefings were updated periodically. 

* Apollo Guidance and Navigation system briefings at the 
Massachusetts Institute of Technology Instrumentation Laboratory 

* Briefings and continuous training on mission photo- 
graphic objectives and use of camera equipment. 

* Extensive pilot participation in reviews of all flight 
procedures for normal as well as emergency situations. 

* Stowage reviews and practice in training sessions in 
the spacecraft, mockups and command module simulators allowed 
the crewmen to evaluate spacecraft stowage of crew-associated 
equipment . 

* More than 400 hours of training per man in command 
module and lunar module simulators at MSC and KSC, including 
closed-loop simulations with flight controllers in the Mission 
Control Center. Other Apollo simulators at various locations 
were used extensively for specialized crew training. 



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-105- 

* Lunar surface briefings and some 20 suited 1-g walk- 
throughs of lunar surface EVA operations covering lunar 
geology and microbiology and deployment of experiments in 
the Apollo Lunar Surface Experiment Package (ALSEP) . Train- 
ing in lunar surface EVA included practice sessions with lunar 
surface sample gathering tools and return containers, cameras, 
the erectable S-band antenna and the modular equipment stowage 
assembly (MESA) housed in the LM descent stage, 

* Proficiency flights in the lunar landing training 
vehicle (LLTV) for the commander. 

* Zero-g and one-sixth g aircraft flights using command 
module and lunar module mockups for EVA and pressure suit 
doffing/donning practice and training. 

* Underwater zero-g training in the MSC Water Immersion 
Facility using spacecraft mockups to further familiarize 
crew with all aspects of CSM-LM docking tunnel intravehicular 
transfer and EVA in pressurized suits. 

* Water egress training conducted in indoor tanks as 
well as in the Gulf of Mexico, included upright ing from the 
Stable II position (apex down) to the Stable I position (apex 
up), egress onto rafts donning Biological Isolation Garments 
(BIGs), decontamination procedures and helicopter pickup. 

* Launch pad egress training from mockups and from the 
actual spacecraft on the launch pad for possible emergencies 
such as fire, contaminants and power failures. 

* The training covered use of Apollo spacecraft fire 
suppression equipment in the cockpit. 

* Planetarium reviews at Morehead Planetarium, Chapel 
Hill, N.C., and at Griffith Planetarium, Los Angeles, Calif., 
of the celestial sphere with special emphasis on the 37 
navigational stars used by the Apollo guidance computer. 



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-106- 
NATIONAL AFSONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, 0. C. 2054* 

BIOGRAPHICAL DATA 

NAME; James Arthur Love 11, Jr. (Captain, U8N) 
NASA Astronaut 

BIRTHPLACE AND DATE: Born March 25, 1928, in Cleveland, Ohio. 
Beach Florida 3 ' BlanChe Lovell » raider. *t Kdgewater 

PHYSICAL DESCRIPTION: Blond hair; blue eyes; height: 5 feet 
11 inches; weight: 170 pounds. 

» 

EDUCATION: Graduated from Juneau High School, Milwaukee, 
Wisconsin; attended the University of Wisconsin for 
d years, then received a Bachelor of Science degree 
from the United States Naval Academy in 1952: presented 
iS ?969! ary D ° Ctorate f ™ m Hlinoi H y weB.Xey2n*uSivS™i?y 

MARITAL STATUS: Married to the former Marilyn Ge-Lach of 

Milwaukee, Wisconsin. Her parents, Mr. and Mrs- Carl 
Cxerlach, are residents of Milwaukee. 

CHILDREN: Barbara L., October 13, 1953; James A., February 
15, 1955; Susan K., July 14, 1958; Jeffrey C, 
January IH , 1966. 

RECREATIONAL INTERESTS: His hobbies are golf, swimming, 
handball, and tennis. fe ' 

ORGANIZATIONS: Member of the Society of Experimental Test 
Pilots and the Explorers Club. 

SPECIAL HONORS: Awarded the NASA Distinguished Service Medal' 
two NASA Exceptional Service Medals, the Navy Astronaut 
^ nSS ^? Navy Dlst ^guished Service Medal, and two 
?S? 21? t i ngUiShed P1 y in g Crosses; recipient of the 

III Zt\ L r a l and G0ld Space Medals (Athens, Greece), 
the American Academy of Achievement Golden Plate Award 
the City of New York Gold Medal in 1969, the City of 
Houston Medal for Valor in 1969, the National Geographic 
Society's Hubbard Medal in 1969, the National Academy 
? ?oJo Vi3i 2 n Arts and Scle nces Special Trustees Award 
m I9b9, and the Institute of Navigation Aw#rd in 1969. 



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-107- 



Co-recipient of the American Astronaut leal Society Flight 
Achievement Awards in 1966 and 1968, the Harmon Inter- 
national Trophy in 1966 and 1967, the Robert H. Goddard 
Memorial Trophy in 196 9, the H. H. Arnold Trophy for 
1969, the General Thomas D. White USAF Space Trophy for 
1968, the Robert J. Collier Trophy for 1968, and the 
1969 Henry G. Bennett Distinguished Service Award. 

EXPERIENCE: Lovell, a Navy Captain, received flight training 
following graduation from Annapolis in 1952. 

He has had numerous naval aviator assignments including 
a 4-year tour as a test pilot at the Naval Air Test 
Center, Patuxent River, Maryland. While there he served 
as program manager for the F4H weapon system evaluation. 
A graduate of the Aviation Safety School of the Univer- 
sity of Southern California, he also served as a flight 
instructor and safety engineer with Fighter Squadron 
101 at the Naval Air Station, Oceana, Virginia. 

He has logged more than 4,407 hours flying time — more than 
3,000 hours in Jet aircraft. 

CURRENT ASSIGNMENT: Captain Lovell was selected as an astronaut 
by NASA in September 1962. He has since served as back- 
up pilot for the Gemini 4 flight and backup command pilot 

for the Gemini 9 flight. 

On December 4, 1965, he and Command pilot Frank Borman were 
launched into space on the history-making Gemini 7 mission. 
The flight lasted 330 hours and 35 minutes, during which 
the following space firsts were accomplished: longest 
manned space flight; first rendezvous of two manned 
maneuverable spacecraft, as Gemini 7 was joined in orbit 
by Gemini 6; and longest multi-manned space flight. It 
was also on this flight that numerous technical and 
medical experiments were completed successfully. 

The Gemini 12 mission, with Lovell and pilot Edwin Aldrin, 
began on November 11, 1966. This 4-day, 59-revolution 
flight brought the Gemini Program to a successful close. 
Major accomplishements of the 94-hour 35-minute flight 
included a third-revolution rendezvous with the previously 
launched Agena (using for the first time backup onboard 
computations due to radar failure); a tethered station- 
keeping exercise,; retrieval of a micrometeorite experl-, 
ment package from the spacecraft exterior; an evaluation 
of the use of body restraints specially designed for 
completing work tasks outside of the spacecraft; and 
completion of numerous photographic experiments, high- 
lights of which are the first pictures taken from space 
of an eclipse of the sun. 

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Gemini 12 ended with retrofire at the beginning of the 
60th revolution, followed by the second consecutive 
fully automatic controlled reentry of a spacecraft, 
and a landing in the Atlantic within 2 1/2 miles of 
the USS WASP. 

As a result of his participation in the Gemini 7 and 12 
flights, Lovell logged 425 hours and 10 minutes In space. 
Aldrin established a new EVA record by completing 5 1/2 
hours outside the spacecraft during two standup EVAs and 
one umbilical EVA. 

Lovell served as command module pilot for the epic six- 
day journey of Apollo 8 — man's maiden voyage to the moon 
—December 21-27, 1968. Apollo 8 was the first manned 
spacecraft to be lifted into near-earth orbit by a 
7 1/2-millIon pound thrust Saturn V launch vehicle, 
and all events in the flight plan occurred as scheduled 
with unbelievable accuracy. 

A "go" for the translunar Injection burn was given midway 
through the second near-earth orbit, and the restart of 
the S-IVB third stage to effect this maneuver increased 
the spacecraft's velocity to place it on an intercept 
course with the moon. Lovell and fellow crew members, 
Prank Borman (spacecraft commander) and William A. 
Anders (lunar module pilot), piloted their spacecraft 
some 223,000 miles to become the first humans to leave 
the earth's influence; and upon reaching the moon on 
December- 24, they performed the first critical maneuver 
to place Apollo 8 into a 60 by 168 nautical miles lunar 
orbit. 

Two revolutions later, the crew executed a second maneuver 
using the spacecraft's 20,500-pound thrust service module 
propulsion system to achieve a circular lunar orbit of 
60 nautical miles. During their ten revolutions of the 
moon, the crew conducted live television transmissions of 
the lunar surface and performed such tasks as landmark 
and Apollo landing site tracking, vertical stereo photo- 
graphy and stereo navigation photography, and sextant 
navigation using lunar landmarks and stars. At the end 
of the tenth lunar orbit, they executed a transearth 
injection burn which placed Apollo 8 on a proper 
trajectory for the return to earth. 



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Th e final leg of the trip required only 58 hours, as 
compared to the 69 hours used to travel to the moon, 
and Apollo 8 came to a successful conclusion on Dec- 
ember 27, 1968. Splashdown occurred at an estimated 
5,000 yeards from the USS Y0RKT0WN, following the 
successful negotiation of a critical 28-mile high 
reentry corridor at speeds close to 25,000 miles per 
hour. 

Captain Lovell has since served as the backup spacecraft 
commander for the Apollo 11 lunar landing mission. He 
has completed three space flights and holds the U.S. 
Astronaut record for time in space with a total of 
572 hours and 10 minutes. 



SPECIAL ASSIGNMENT: In addition to his regular duties as an 

astronaut, Captain Lovell continues to serve as Special 
Consultant to the President's Council on Physical 
Fitness and Sports — an assignment he has held since 
June 1967. 

CURRENT SALARY: $1,717-2 8 per month. 



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NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, D. C. 20546 

BIOGRAPHICAL DATA 

NAME: Thomas Kenneth Mattingly II (Lieutenant Commander, USN) 
NASA Astronaut 

BIRTHPLACE AND DATE: Born In Chicago, 111., March 17, 1936. 
His parents, Mr. and Mrs. Thomas K. Mattingly, now 
reside in Hialeah, Fla. 

PHYSICAL DESCRIPTION: Brown hair; blue eyes; height: 5 feet 
10 inches; weight: 140 pounds. 

EDUCATION: Attended Florida elementary and secondary 

schools and is a graduate of Miami Edison High School, 
Miami, Fla.; received a Bachelor of Science degree in 
Aeronautical Engineering from Auburn University in 
1958. 

MARITAL STATUS: Single 

RECREATIONAL INTERESTS: Enjoys water skiing and playing 
handball and tennis . 

ORGANIZATIONS: Member of the American Institute of Aero- 
nautics and Astronautics and the U.S. Naval Institute. 

EXPERIENCE: Prior to reporting for duty at the Manned 

Spacecraft Center, he was a student at the Air Force 
Aerospace Research Pilot School. 

He began his Naval career as an Ensign in 1958 and 
received his wings in I960. He was then assigned to 
VA-35 and flew A1H aircraft aboard the USS SARATOGA 
from I960 to 1963 . In July 1963, he served in VAH-11 
deployed aboard the USS FRANKLIN D. ROOSEVELT where 
he flew the A3B aircraft for two years. 

He has logged 3,700 hours of flight time — 1,946 hours 
in jet aircraft. 

CURRENT ASSIGNMENT: Lt Commander Mattingly is one of the 19 

astronauts selected by NASA In April 1966. He served 

as a member of the astronaut support crews for the 
Apollo 8 and 11 missions. 

CURRENT SALARY: $1,293-33 per month. 



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NATIONAl AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, D. C. 20546 

BIOGRAPHICAL DATA 

NAME: Fred Wallace Haise, Jr. (Mr.) 
NASA Astronaut 

BIRTHPLACE AND DATE: Born In Biloxi, Miss., on Nov. 14, 1933; 
his mother, Mrs. Fred W. Haise, Sr., resides in Biloxi. 

PHYSICAL DESCRIPTION: Brown hair; brown eyes; height: 5 feet 
9 1/2 inches; weight: 150 pounds. 

EDUCATION: Graduated from Biloxi High School, Biloxi, Miss.; 

attended Perkinston Junior College (Association of Arts); 
received a Bachelor of Science degree with honors in 
Aeronautical Engineering from the University of Oklahoma 
in 1959. 

MARITAL STATUS: Married to the former Mary Griffin Grant of 
Biloxi, Miss. Her parents, Mr. and Mrs. William J. 
Grant, Jr., reside in Biloxi. 

CHILDREN: Mary M. , January 25, 1956; Frederick T., May 13, 1958; 
Stephen W., June 30, 1961. 

ORGANIZATIONS: Member of the Society of Experimental Test 
Pilots, Tau Beta Pi, Sigma Gamma Tau, and Phi Theta 
Kappa. 

SPECIAL HONORS: Recipient of the A. B. Honts Trophy as the 
outstanding graduate of class 64a from the Aerospace 
Research Pilot School in 1964; awarded the American 
Defense Ribbon and the Society of Experimental Test 
Pilots Ray E. Tenhoff Award for 1966. 

EXPERIENCE: Halse was a research pilot at the NASA Flight 
Research Center at Edwards, Calif., before coming to 
Houston and the Manned Spacecraft Center; and from 
September 1959 to March 1963> he was a research pilot 
at the NASA Lewis Research Center in Cleveland, Ohio. 
During this time, he authored the following papers which 
have been published: a NASA TND , entitled "An Evaluation 
of the Flying Qualities of Seven General-Aviation Air- 
craft ;" NASA TND 3380, "Use of Aircraft for Zero Gravity 
Environment, May 1966;" SAE Business Aircraft Conference 
Paper, entitled "An Evaluation of General-Aviation Air- 
craft Flying Qualities," March 30-April 1, 1966; and a 
paper delivered at the tenth symposium of the Society of 

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Experlmental Test Pilots, entitled "A Quantitative/ 
Qualitative Handling Qualities Evaluation of Seven 
General-Aviation Aircraft , " 1966, 

He was the Aerospace Research Pilots School's out- 
standing graduate of Class 6 MA and served with the 
U.S. Air Force from October 1961 to August 1962 as a 
tactical fighter pilot and as Chief of the l6kth 
Standardization-Evaluation Plight of the 164th Tactical 
Fighter Squadron at Mansfield, Ohio. From March 1957 
to September 1959 , he was a fighter-interceptor pilot 
with the 185th Fighter Interceptor Squadron in the Okla- 
homa Air National Guard. 

He also served as a tactics and all weather flight 
instructor in the U.S. Navy Advanced Training Command 
at NAAS Kingsville, Texas, and was assigned as a U.S. 
Marine Corps fighter pilot to VMF-533 and 114 at MCAS 
Cherry Point, N.C., from March 195^ to September 1956. 

His military career began in October 1952 as a Naval 
Aviation Cadet at the Naval Air Station in Pensacola, 
Fla. 

He has accumulated 5,800 hours flying time, including 
3,000 hours in jets. 

CURRENT ASSIGNMENT: Mr. Haise is one of the 19 astronauts 
selected by NASA in April 1966. He served as backup 
lunar module pilot for the Apollo 8 and 11 missions. 



CURRENT SALARY: $1,698.00 per month. 



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LAUNCH COMPLEX 39 



Launch Complex 39 facilities at the Kennedy Space 
Center were planned and built specifically for the Apollo 
Saturn V, the space vehicle being used in the United States 
manned lunar exploration program. 

Complex 39 introduced the mobile concept of launch 
operations in which the space vehicle is thoroughly checked 
out in an enclosed building before it is moved to the launch 
pad for final preparations. This affords greater protection 
from the elements and permits a high launch rate since pad 
time is minimal. 

Saturn V stages are shipped to the Kennedy Space 
Center by ocean-going vessels and specially designed air- 
craft. Apollo spacecraft modules are transported by air 
and first taken to the Manned Spacecraft Operations Building 
in the Industrial Area south of Complex 39 for preliminary 
checkout, altitude chamber testing, and assembly. 

Apollo 12 is the sixth Saturn V/Apollo space vehicle 
to be launched from Complex 39* s Pad A, one of two octagonal 
launch pads which are 3,000 feet across. The major components 
of Complex 39 include: 

1. T he Vehicle Assembly Building , heart of the complex, 
is where the 363-foot-tall space vehicle is assembled and 
tested. It contains 129-5 million cubic feet of space, covers 
eight acres, is 716 feet long and 518 feet wide. Its high 
bay area, 525 feet high, contains four assembly and checkout 
bays and its low bay area - 210 feet high, 442 feet wide and 
274 feet long - contains eight stage-preparation and check- 
out cells. There are 141 lifting devices in the building, 
ranging from one-ton hoists to two 250-ton high lift bridge 
cranes . 

2. The Launch Control Center , a four-story structure 
adjacent and to the south of the Vehicle Assembly Building 
is a radical departure from the dome-shaped, "hardened" 
blockhouse at older launch sites. The Launch Control Center 
is the electronic "brain" of Complex 39 and was used for 
checkout and test operations while Apollo 12 was being as- 
sembled inside the Vehicle Assembly Building high bay. Three 
of the four firing rooms contain identical sets of control 
and monitoring equipment so that launch of one vehicle and 
checkout of others may continue simultaneously. Each firing 
room is associated with a ground computer facility to provide 
data links with the launch vehicle on Its mobile launcher at 
the pad or inside the Vehicle Assembly Building. 

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3* The Mobile Launcher , 445 feet tall and weighing 
12 million pounds, is a transportable launch base and um- 
bilical tower. for the space vehicle. 

4. The Transporters , used to move mobile launchers 
into the Vehicle Assembly Building and then - with their 
space vehicles - to the launch pad, weigh six million pounds 
and are among the largest tracked vehicles known. The Trans- 
porters - there are two - are 131 feet long and 114 feet wide. 
Powered by electric motors driven by two 2 3 750-horsepower 
diesel engines, the vehicles move on four double-tracked 
crawlers, each 10 feet high and 40 feet long. Maximum speed 
is about one-mile-per-hour loaded and two miles-per-hour 
unloaded. The three and one-half mile trip to Pad A with a 
mobile launcher and space vehicle takes approximately seven 
hours. Apollo 12 rollout to the pad occurred on December 
15, 1969. <-enu>er 

.5. The Crawlerway is the roadway for the transporter 
and is 131 feet wide divided by a median strip. This is the 
approximate width of an eight-lane turnpike and the roadbed 
is designed to accommodate a combined weight of more than 18 
million pounds. 

6. The Mobile Service Structure is a 402-foot-tall, 9.8 
million pound tower used to service the Apollo space vehicle 
at the pad. Moved into place about the Saturn V/Apollo space 
vehicle and its mobile launcher by a transporter, it contains 
five work platforms and provides 360-degree platform access to 
the vehicle being prepared for launch. It is removed to a 
parking area about 11 hours before launch. 

7. A Water Deluge System will provide about a million 
gallons of industrial water for cooling and fire prevention 
during the launch of Apollo 13. The water is used to cool 
the mobile launcher, the flame trench and the flame deflector 
above which the mobile launcher is positioned. 

8- The Flame Deflector is an "A"-shaped, 1.3 million 
pound structure moved into the flame trench beneath the launcher 
prior to launch. It is covered with a refractory material 
designed to withstand the launch environment , The flame trench 
itself is 58 feet wide and approximately six feet above mean 
sea level at the base. 



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9. The Pad Areas - A and B - are octagonal in shape 
and have center hardstands constructed of heavily reinforced 
concrete. The top of Pad A stands about 48 feet above sea 
level. Saturn V propellants - liquid oxygen, liquid hydrogen 
and RP-1, the latter a high grade kerosene - are stored in 
large tanks spaced near the pad perimeter and carried by pipe- 
lines from the tanks to the pad, up the mobile launcher and into 
the launch vehicle propellant tanks. Also located in the pad 
area are pneumatic, high pressure gas, electrical, and industrial 
water support facilities. Pad B, used for the launch of Apollo 
10, is located 8,700 feet north of Pad A. 



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MIS5I0N CONTROL CENTER 

The Mission Control Center at the Manned Spacecraft 
Center, Houston, is the focal point for Apollo flight control 
activities. The center receives tracking and telemetry data 
from the Manned Space Plight Network which in turn is pro- 
cessed by the MCC Real-Time Computer Complex for display to 
flight controllers in the Mission Operations Control Room 
(MOCR) and adjacent staff support rooms. 

Console positions in the two identical MOCRs in 
Mission Control Center fall into three basic operations 
groups: mission command and control, systems operations, 
and flight dynamics. 

Positions in the command and control group are: 

* Mission Director — responsible for overall mission 
conduct . 

* Flight Operations Director — represents MSC management. 

* Flight Director — responsible for operational decisions 
and actions in the MOCR. 

* Assistant Flight Director — assists flight director 
and acts in his absence. 

* Flight Activities Officer — develops and coordinates 
flight plan. 

* Department of Defense Representative — coordinates 
and directs DOD mission support. 



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* Network Controller — responsible to FD for Manned 
Space Flight Network status and troubleshooting; MCC equip- 
ment operation. 

* Surgeon — monitors crew medical condition and informs 
FD of any medical situation affecting mission. 

* Spacecraft Communicator (Capcom) — serves as voice 
contact with flight crew. 

* Experiments Officer — coordinates operation and 
control of onboard flight experiments. 

* Public Affairs Officer — reports mission progress 

to public through commentary and relay of live air-to-ground 
transmissions . 

Systems Operations Group: 

* Environmental, Electrical and Instrumentation 
Engineer (EECOM) — monitors and troubleshoots command/service 
module environmental, electrical, and sequential systems. 

* Guidance, Navigation and Control Engineer (GNC) — 
monitors and troubleshoots CSM guidance, navigation, control, 
and propulsion systems. 

* LM Environmental and Electrical Engineer (TELCOM) — 
LM counterpart to EECOM. 

* LM Guidance, Navigation and Control Engineer (Control)-- 
LM counterpart to GNC. 

* Booster Systems Engineer (BSE) (three positions) — 
responsible for monitoring launch vehicle nerformance and for 
sending function commands. 

* Communications Systems Engineer (CSE) (call sign INCO) 
and Operations and Procedures Officer (O&P) — share respon- 
sibility for monitoring and troubleshooting spacecraft and 
lunar surface communication systems and for coordinating MCC 
procedures with other NASA centers and the network. 

Flight Dynamics Group: 

• Flight Dynamics Officer (FIDO) — monitors powered 
flight events and plans spacecraft maneuvers. 

* Retrofire Officer (Retro) — responsible for plurirn njr 
deorbit maneuvers in Earth orbit and entry calculations on 
lunar return trajectories. 



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* Guidance Officer (Guldo) — responsible for monitoring 
and updating CSM and LM guidance systems and for monitoring 
systems performance during powered flight. 

Each MOCR operations group has a staff support room 
on the same floor in which detailed monitoring and analysis 
is conducted. Other supporting MCC areas include the space- 
flight Meteorological Room, the Space Environment (radiation) 
Console, Spacecraft Planning and Analysis (SPAN) Room for 
detailed spacecraft performance analysis, Recovery Operations 
Control Room and the Apollo Lunar Surface Experiment Package 
Support Room. 

Located on the first floor of the MCC are the communica- 
tions, command, and telemetry system (CCATS) for processing 
incoming data from the tracking network, and the real-time 
computer complex (RTCC) which converts flight data into dis- 
plays useable to MOCR flight controllers. 



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MANNED SPACE FLIGHT NETWORK 

The worldwide Manned Space Plight Network (MSFN) 
provides reliable, continuous, and instantaneous com- 
munications with the astronauts, launch vehicle, and 
spacecraft from liftoff to splashdown. Following the 
flight, the network will continue in support of the 
link between Earth and the Apollo experiments left on the 
lunar surface by the Apollo crew. 

The MSFN is maintained and operated by the NASA 
Goddard Space Flight Center, Greenbelt, Md., under the 
direction of NASA's Office of Tracking and Data Acquisition. 
In the MSFN Operations Center (MSFNOC) at Goddard, the 
Network Director and his team of Operations Managers, with the 
assistance of a Network Support Team, keep the entire complex 
tuned for the mission support. Should Houston's mission con- 
trol center be seriously impaired for an extended time, the 
Goddard Center becomes an emergency mission control center. 

The MSFN employs 12 ground tracking stations equipped 
with 30- and 85-foot antennas, an instrumented tracking ship, 
and four instrumented aircraft. For Apollo 13 > the network 
will be augmented by the 210-foot antenna systems at Goldstone, 
Calif, and at Parkes, Australia, (Australian Commonwealth 
Scientific and Industrial Research Organization) . 

NASA Communications Network (NASCOM) . The tracking 
network is linked together by the NASA Communications Network. 
All information flows to and from MCC Houston and the Apollo 
spacecraft over this communications system. 

The NASCOM consists of almost three million circuit 
miles of diversely routed communications channels. It uses 
satellites, submarine cables, land lines, microwave systems, 
and high frequency radio facilities for access links. 

NASCOM control center is located at Goddard. Regional 
communication switching centers are In London, Madrid, Can- 
berra, Australia; Honolulu and Guam. 



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TRACKING TH 

GOLDSTONE, CALIFORNIA 




MOON 



MADRID, SPAIN 




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Three Intelsat communications satellites will be used 
for Apollo 13. One satellite over the Atlantic will link 
Goddard with stations at Madrid, Canary Islands, Ascension 
and the Vanguard tracking ship. Another Atlantic satellite 
will provide a direct link between Madrid and Goddard for 
TV signals received from the spacecraft. The third 
satellite over the mid-Pacific will link Carnarvon, Canberra, 
and Hawaii with Goddard through a ground station at Brewster 
Plats, Wash. 

At Goddard, NASCOM switching computers simultaneously 
send the voice signals directly to the Houston flight controllers 
and the tracking and telemetry data to computer processing 
complexes at Houston and Goddard. The Goddard Real Time Com- 
puting Complex verifies performance of the tracking network and 
uses the collected tracking data to drive displays in the 
Goddard Operations Control Center. 

Establishing the Link — The Merritt Island tracking 
station monitors prelaunch test, the terminal countdown, and 
the first minutes of launch. 

An Apollo instrumentation ship (USNS VANGUARD) fills 
the gaps beyond the range of land tracking stations. For 
Apollo 13 this ship will be stationed in the Atlantic to 
cover the Insertion into Earth orbit . Apollo Instrumented 
aircraft provide communications support to the land tracking 
stations during translunar injection and reentry and cover 
a selected abort area in the event of "no-go" decision after 
insertion into Earth orbit. 

Lunar Bound - Approximately one hour after the space- 
craft has been Injected into its translunar trajectory (some 
10,000 miles from the Earth), three prime tracking stations 
spaced nearly equidistant around the Earth will take over 
tracking and communicating with Apollo. 

Each of the prime stations, located at Golds tone, 
Madrid and Canberra, has a dual system for use when tracking 
the command module in lunar orbit and the lunar module in 
separate flight paths or at rest on the Moon. These stations 
are equipped with 85-foot antennas . 

The Return Trip — To make an accurate reentry, data 
from the tracking stations are fed into the MCC computers 
to develop necessary information for the Apollo 13 crew. 



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Appropriate MSFN stations, Including the aircraft In the 
Pacific, provide support during the reentry. 

Through the journey to the Moon and return, television 
will be received from the spacecraft at the three prime 
stations. In addition, a 210-foot antenna at Golds tone (an 
antenna of NASA's Deep Space Network) will augment the 
television coverage while Apollo 13 is near and on the Moon. 
For black and white TV, scan converters at the stations per- 
mit immediate transmission of commercial quality TV via NASCOM 
to Houston, where it will be released to U.S. TV networks. 

Black and white TV can be released simultaneously in 
Europe and the Far East through the MSFN stations in Spain 
and Australia. 

For color TV, the signal will be converted to commercial 
quality at the MSC Houston. A black and white version of the 
color signal can be released locally simultaneously through 
the stations in Spain and Australia. 

Network Computers 

At fraction-of-a-second intervals, the network's digital 
data processing systems, with NASA's Manned Spacecraft Center 
as the focal point, "talk" to each other or to the spacecraft. 
High-speed computers at the remote sites (tracking ship in- 
cluded) relay commands or "up-link" data on such matters as 
control of cabin pressure, orbital guidance commands, or "go- 
no-go" indications to perform certain functions. 

When information originates from Houston, the computers 
refer to their pre-programmed information for validity before 
transmitting the required data to the spacecraft. 

Such "up-link" information is communicated at a rate of 
about 1,200 bits-per-second. Communication of spacecraft data 
between remote ground sites and the Mission Control Center, 
via high-speed communications links, occurs at twice the rate. 
Houston reads information from these ground sites at 8,800 
bits-per-second. 

The computer systems perform many other functions, including: 

Assuring the quality of the transmission lines by con- 
tinually testing data paths. 

Verifying accuracy of the messages. 

Constantly updating the flight status. 



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For "down- link" data, sensors built into the spacecraft 
continually sample cabin temperature, pressure, and physical 
information on the astronauts such as heartbeat and respira- 
tion. These data are transmitted to the ground stations at 
51.2 kilobits (12,800 decimal digits) per second. 

At MCC the computers : 

Detect and select changes or deviations, compare with 
their stored programs, and indicate the problem areas or 
pertinent data to the flight controllers; 

Provide displays to mission personnel; 

Assemble output data in proper formats; 

Log data on magnetic tape for the flight 
controllers. 

The Apollo Ship Vanguard 

The USNS Vanguard will perform tracking, telemetry, 
and communication functions for the launch phase and Earth 
orbit insertion. Vanguard will be stationed about 1,000 miles 
southeast of Bermuda (28 degrees N. s 49 degrees W.). 

Apollo Range Instrumentation Aircraft (ARIA) 

During the Apollo 13 TLI maneuver, two ARIA will record 
telemetry data from Apollo and relay voice communication 
between the astronauts and the Mission Control Center at 
Houston. The ARIA will be located between Australia and 
Hawaii . 

For reentry, two ARIA will be deployed to the landing 
area to relay communications between Apollo and Mission Con- 
trol at Houston and provide position information on the space- 
craft after the blackout phase of reentry has passed. 

The total ARIA fleet for Apollo missions consists of 
four EC-135A (Boeing 707) jets with 7-foot parabolic antennas 
installed in the nose section. 



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CONTAMINATION CONTROL PROGRAM 



In 1966 an Interagency Committee on Back Contamination 
(ICBC) was established to assist NASA in developing a pro- 
gram to prevent contamination of the Earth from lunar mat- 
erials following manned lunar exploration and to review and 
approve plans and procedures to prevent back contamination. 
Committee membership includes representatives from Public 
Health Service, Department of Agriculture, Department of the 
Interior, NASA, and the National Academy of Sciences. 

The Apollo Back Contamination Program can be divided 
into three phases. The first phase covers procedures which 
are followed by the crew while in flight to reduce and, if 
possible, eliminate the return of lunar surface contaminations 
in the command module . 

The second phase includes recovery, isolation, and 
transport of the crew, spacecraft, and lunar samples to the 
Manned Spacecraft Center. The third phase encompasses 
quarantine operations and preliminary sample analysis in the 
Lunar Receiving Laboratory. 

A primary step in preventing back contamination is 
careful attention to spacecraft cleanliness following lunar 
surface operations. This includes use of special cleaning 
equipment, stowage provisions for lunar-exposed equipment, 
and crew procedures for proper "housekeeping. " 

Prior to reentering the LM after lunar surface explora- 
tion, the crewmen brush lunar surface dust or dirt from the 
space suit using special brushes. They will scrape their 
overboots on the LM footpad and while ascending the LM ladder, 
dislodge any clinging particles by a kicking action. 

After entering and pressurizing the LM cabin, the crew 
doff their portable life support system, oxygen purge system, 
lunar boots, EVA gloves, etc. 

Following LM rendezvous and docking with the CM, the CM 
tunnel will be pressurized and checks made to insure that an 
adequate pressurized seal has been made. During the period, 
some of the equipment may be vacuumed. 



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The lunar module cabin atmosphere will be circulated 
through the environmental control system suit circuit lithium 
hydroxide (LiOH) canister to filter particles from the atmosphere 
A minimum of five hours weightless operation and filtering will 
essentially eliminate the original airborne particles . 

The CM pilot will transfer lunar surface equipment stowage 
bags into the LM one at a time. The equipment transferred will 
be bagged before being transferred. The only equipment which 
will not be bagged at this time are the crewmen r s space suits 
and flight logs . 

Command Module Operations - Through the use of operational 
and housekeeping procedures the command module cabin will be 
purged of lunar surface and/or other particulate contamination 
prior to Earth reentry. These procedures start while the LM 
is docked with the CM and continue through reentry into the 
Earth f s atmosphere . 

During subsequent lunar orbital flight and the transearth 
phase, the command module atmosphere will be continually 
filtered through the environmental control system lithium 
hydroxide canister. This will remove essentially all airborne 
dust particles. After about 96 hours operation essentially 
none of the original contaminates will remain. 

Lunar Mission Recovery Operations 

Following landing and the attachment of the flotation 
collar to the command module, a swimmer will open the space- 
craft hatch, pass in three clean flight coveralls and three 
filter masks and close the hatch. 

Crew retrieval will be accomplished by helicopter to 
the carrier and subsequent crew transfer to the Mobile 
Quarantine Facility. The spacecraft will be retrieved by the 
aircraft carrier and Isolated. 



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LUNAR RECEIVING LABORATORY (LRL) 

The final phase of the back contamination program is 
completed in the MSC Lunar Receiving Laboratory. The crew 
and spacecraft are quarantined for a minimum of 21 days after 
completion of lunar EVA operations and are released based upon 
the completion of prescribed test requirements and results. 
The lunar sample will be quarantined for a period of 50 to 80 
days depending upon results of extensive biological tests. 

The LRL serves four basic purposes: 

Quarantine of crew and spacecraft, the containment 
of lunar and lunar-exposed materials, and quarantine testing 
to search for adverse effects of lunar material upon terrestrial 
life. 

The preservation and protection of the lunar samples. 

The performance of time critical investigations. 

The preliminary examination of returned samples to 
assist in an intelligent distribution of samples to principal 
investigators. 

The LRL has the only vacuum system in the world with 
space gloves operated by a man leading directly into a vacuum 
chamber at pressures of about 10 billionth of an atmosphere. 
It has a low level counting facility, the background count is 
an order of magnitude better than other known counters. 
Additionally, it is a facility that can handle a large variety 
of biological specimens inside Class III biological cabinets 
designed to contain extremely hazardous pathogenic material. 

The LRL covers 83,000 square feet of floor space and 
includes a Crew Reception Area (CRA), Vacuum Laboratory, 
Sample Laboratories (Physical and Bio-Science) and an 
administrative and support area. Special building systems 
are employed to maintain air flow into sample handling areas 
and the CRA, to sterilize liquid waste, and to incinerate 
contaminated air from the primary containment systems. 

The biomedical laboratories provide for quarantine tests 
to determine the effect of lunar samples on terrestrial life. 
These tests are designed to provide data upon which to base 
the decision to release lunar material from quarantine. 



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Araong the tests : 

a. Lunar material will be applied to 12 different 
culture media and maintained under several environmental 
conditions . The media will be observed for bacterial or 
fungal growth. Detailed inventories of the microbial flora 
of the spacecraft and crew have been maintained so that any 
living material found in the sample testing can be compared 
against this list of potential contaminants taken to the Moon 
by the crew or spacecraft. 

b. Six types of human and animal tissue culture cell 
lines will be maintained in the laboratory and together with 
embryonated eggs are exposed to the lunar material. Based on 
cellular and/or other changes, the presence of viral material 
can be established so that special tests can be conducted to 
identify and isolate the type of virus present. 

c. Thirty-three species of plants and seedlings will 
be exposed to lunar material. Seed germination, growth of 
plant cells or the health of seedlings are then observed, 
and histological, microbiological and biochemical techniques 
are used to determine the cause of any suspected abnormality. 

d. A number of lower animals will be exposed to lunar 
material, including germ- free mice, fish, birds, oysters, 
shrimp, cockroaches, houseflies, planaria, paramecia and 
euglena. If abnormalities are noted, further tests will be 
conducted to determine if the condition is transmissible from 
one group to another. 

The crew reception area provides biological containment 
for the flight crew and 12 support personnel. The nominal 
occupancy is about 14 days but the facility is designed and 
equipped to operate for considerably longer. 

Sterilization and Release of the Spacecraft 

Postflight testing and inspection of the spacecraft is 
presently limited to investigaiton of anomalies which happened 
during the flight. Generally, this entails some specific 
testing of the spacecraft and removal of certain components of 
systems for further analysis. The timing of postflight testing 
is important so that corrective action may be taken for sub- 
sequent flights . 



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The schedule calls for the spacecraft to be returned 
to port where a team will deactivate pyrotechnics, and flush 
and drain fluid systems (except water) . This operation will 
be confined to the exterior of the spacecraft. The spacecraft 
will then be flown to the LRL and placed in a special room for 
storage, sterilization, and postf light checkout. 



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LUNAR RECEIVING LABORATORY TENTATIVE SCHEDULE 

April 20 Activate secondary barrier; support people enter 
Crew Reception Area and Central Status Station 
manned; LRL on mission status. 

April 21 Command module landing, recovery. 

April 22 First sample return container (SRC) arrives. 

April 23 First SRC opened in vacuum lab, second SRC arrives; 
film, tapes, LM tape recorder begin decontamination; 
second SRC opened in Bioprep lab. 

Anril 24 First sample to Radiation Counting Laboratory. 

April 26 Core tube moves from vacuum lab to Physical- 
Chemical Lab. 

April 26 MQP arrives; contingency sample goes to Physical- 
Chemical Lab; rock description begun in vacuum lab. 

April 27 Biosample rocks move from vacuum lab to Bioprep 
"Lab; core tube prepared for biosample. 

April 28 Spacecraft arrives. 

April 29 Biosample compounded, thin-section chips sterilized 
out to Thin-Section Lab, remaining samples from 
Bioprep Lab canned. 

May 1 Thin-section preparation complete, biosample prep 
complete, transfer to Physical-Chemical Lab 
complete, Bioprep Lab cleanup complete. 

May 3 Biological protocols, Physical-Chemical Lab rock 
description begin. 

May 8 Crew released from CRA 

May 26 Rock description complete, Preliminary Examination 
Team data from Radiation Counting Lab and Gas 
Analysis Lab complete . 

May 28 PET data write-up and sample catalog preparation 
begin. 



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May 30 



June 


1 


June 


2 


June 


6 


June 


8 


June 


10 


June 


1H 



Data summary for Lunar Sample Analysis Planning 
Team (LSAPT) complete. 

LSAPT arrives. 

LSAPT briefed on PET data, sample packaging begins. 

Sample distribution plan complete, first batch 
monopole samples canned. 

Monopole experiment begins . 

Initial release of Apollo 13 samples; spacecraft release 

Spacecraft equipment released 



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SCHEDULE FOR TRANSPORT OF SAMPLES, SPACECRAFT AND CREW 

Samples 

The first Apollo 13 sample return container (SRC) will be 
flown by helicopter from the deck of the USS Iwo Jima to 
Christmas Island, from where it will be flown by C-130 aircraft 
to Hawaii. The SRC, half the mission onboard film and any 
medical samples ready at the time of helicopter departure from 
the Iwo Jima * will be transferred to an ARIA (Apollo Range 
Instrumented Aircraft) at Hawaii for the flight to Ellington 
APB, six miles north of the Manned Spacecraft Center, with an 
estimated time of arrival at 11:30 a.m. EST April 22. 

The second SRC and remainder of onboard film and medical 
samples will follow a similar sequence of flights the following 
day and will arrive at Ellington AFB at an estimated time of 1 
am EST April 23. The SRCs will be moved by auto from Ellington 
AFB to the Lunar Receiving Laboratory. 

Spacecraft 

The spacecraft should be aboard the Iwo Jima about two 
hours after crew recovery. The ship will arrive in Hawaii 
at 2 pm EST April 25 and the spacecraft will be offloaded and 
transferred after deactiviation to an aircraft for the flight 
to Ellington AFB, arriving April 28. The spacecraft will be 
trucked to the Lunar Receiving Laboratory where it will enter 
quarantine . 

Crew 

The flight crew is expected to enter the Mobile Quarantine 
Facility (MQF) on the Iwo Jima about 90 minutes after splash- 
down. Upon arrival at Hawaii, the MQF will be offloaded and 
placed aboard a C-l4l aircraft for the flight to Ellington AFB, 
arriving at 1 am EST April 25. A transporter truck will move 
the MQF from Ellington AFB to the Lunar Receiving Laboratory — 
about a two-hour trip. 



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APOLLO PROGRAM MANAGEMENT 

The Apollo Program is the responsibility of the Office 
of Manned Space Flight (OMSF), National Aeronautics and Space 
Administration, Washington, D. C. Dale D. Myers is Associate 
Administrator for Manned Space Plight. 

NASA Manned Spacecraft Center (MSC), Houston, is responsible 
for development of the Apollo spacecraft, flight crew training 
and flight control. Dr. Robert R. Gilruth is Center Director.' 

NASA Marshall Space Flight Center (MSFC), Huntsville, Ala. 
is responsible for development of the Saturn launch vehicles. ' 
Dr. Eberhard Rees is Center Director. 

NASA John F. Kennedy Space Center (KSC), Fla. , is responsible 
for Apollo/Saturn launch operations. Dr. Kurt H. Debus is 
Center Director. 

The NASA Office of Tracking and Data Acquisition (OTDA) 
directs the program of tracking and data flow on Apollo. 
Gerald M. Truszynski is Associate Administrator for Tracking 
and Data Acquisition. 

NASA Goddard Space Flight Center (GSFC), Greenbelt, Md., 
manages the Manned Space Flight Network and Communications 
Network. Dr. John F. Clark is Center Director. 

The Department of Defense is supporting NASA in Apollo 13 
during launch, tracking and recovery operations. The Air Force 
Eastern Test Range is responsible for range activities during 
launch and down-range tracking. Recovery operations include the 
use of recovery ships and Navy and Air Force aircraft. 



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Apollo/Saturn Officials 
NASA Headquarters 

Dr. Rocco A. Petrone Apollo Program Director, OMSF 

Chester M. Lee (Capt., USN, Ret.) Apollo Mission Director, OMSF 
Col. Thomas H. McMullen (USAF) 



John D. Stevenson (Maj . Gen., 
USAF, Ret.) 



Apollo Assistant Mission 
Director, OMSF 

Director of Mission Operations, 
OMSF 



Maj. Gen. James W. Humphreys, Jr. Director of Space Medicine, OMSF 
(USAF, MC) 

John K. Holcomb,(Capt. , USN, Ret.) Director of Apollo 

Operations, OMSF 

Lee R. Scherer, (Capt . , USN, Ret.) Director of Apollo Lunar 

Exploration, OMSF 



James C. Bavely 

Marshall Space Flight Center 
Lee B. James 
Dr. F. A. Speer 

Roy E. Godfrey 
Matthew W. Urlaub 

William F. LaHatte 

Charles H. Meyers 

Frederich Duerr 

William D. Brown 



Chief of Network Operations 
Branch, OTDA 



Director, Program Management 

Manager, Mission Operations 
Office 

Manager, Saturn Program Office 

Manager, S-IC Stage, Saturn 
Program Office 

Manager, S-II Stage, Saturn 
Program Office 

Manager (Acting), S-IVB Stage, 
Saturn Program Office 

Manager, Instrument Unit, 
Saturn Program Office 

Manager, Engine Program Office 



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Kennedy Space Center 
Walter J. Kapryan 
Raymond L. Clark 
Edward R. Mathews 
Dr . Hans F . Gruene 

John J. Williams 

Paul C. Donnelly 

Isom A. Rigell 

Manned Spacecraft Center 

Col.. James A. McDivitt, (USAF) 

Donald K. Slayton 

Sigurd A. Sjoberg 
Milton L. Windier 
Gerald Griffin 
Glynn S. Lunney 
Eugene F. Kranz 
Dr. Charles A. Berry 

Goddard Space Flight Center 
Ozro M„ Covington 

William P. Varson 

H. William Wood 

Tecwyn Roberts 

L. R. Stelter 



Director of Launch Operations 

Director of Technical Support 

Apollo Program Manager 

Director, Launch Vehicle 
Operations 

Director, Spacecraft Operations 

Launch Operations Manager 

Deputy Director for Engineering 

Manager, Apollo Spacecraft 
Program 

Director, Flight Crew 
Operations 

Director, Flight Operations 

Flight Director 

Flight Director 

Flight Director 

Flight Director 

Director, Medical Research 
and Operations 



Director of Manned Flight 
Support 

Chief, Manned Flight Planning 
& Analysis Division 

Chief, Manned Flight Operations 
Division 

Chief, Manned Flight Engineering 
Division 



Chief, NASA Communications 
Division. 



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Department of Defense 



Maj. Gen. David M. Jones, (USAF) 



Rear Adm. Wra. S. Guest, (USN) 



Rear Adm. Donald C. Davis, (USN) 



Col. Kenneth J. Mask, (USAF) 



MaJ. Gen. Allison C. Brooks, 
(USAF) 



DOD Manager of Manned Space 
Flight Support Operations, 
Commander of USAF Eastern 
Test Range 

Deputy DOD Manager of Manned 
Space Flight Support Operations, 
Commander Task Force 140, 
Atlantic Recovery Area 

Commander Task Force 130, 
Pacific Recovery Area 

Director of DOD Manned Space 
Flight Support Office 

Commander Aerospace Rescue and 
Recovery Service 



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Major Apollo/Saturn V Contractors 



Contractor 



Item 



Bellcomm 
Washington, D. C. 

The Boeing Co. 
Washington, D. C. 

General Electric-Apollo Systems 
Daytona Beach, Pla 

North American Rockwell Corp. 
Space Div. , Downey, Calif. 

Grumman Aircraft Engineering 
Corp., Bethpage, N.Y. 

Massachusetts Institute of 
Technology, Cambridge, Mass. 

General Motors Corp., AC 
Electronics Div., Milwaukee, Wis. 

TRW Inc. 
Systems Group 
Redondo Beach, Calif. 

Avco Corp., Space Systems 
Div., Lowell, Mass. 

North American Rockwell Corp. 
Rocket dyne Div. 
Canoga Park, Calif. 

The Boeing Co. 
New Orleans, 



North American Rockwell Corp. 

Space Div. 

Seal Beach, Calif, 

McDonnell Douglas Astronautics 
Co., Huntington Beach, Calif. 



Apollo Systems Engineering 



Technical Integration and 
Evaluation 

Apollo Checkout, and Quality and 
Reliability 

Command and Service Modules 



Lunar Module 



Guidance & Navigation 

(Technical Management) 

Guidance & Navigation 
(Manufacturing) 

Trajectory Analysis 

LM Descent Engine 

LM Abort Guidance System 

Heat Shield Ablative Material 



J-2 Engines, F-l Engines 



First Stage (SIC) of Saturn V 
Launch Vehicles, Saturn V 
Systems Engineering and Inte- 
gration, Ground Support Equip- 
ment 

Development and Production of 
Saturn V Second Stage (S-II) 



Development and Production of 
Saturn V Third Stage (S-IVB) 



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International Business Machines 
Federal Systems Div. 
Huntsville, Ala. 

Bendix Corp. 

Navigation and Control Div. 

Teterboro, N.J. 

Federal Electric Corp. 



Bendix Field Engineering Corp 



Catalytic-Dow 



Hamilton Standard Division 
United Aircraft Corp. 
Windsor Locks, Conn. 

ILC Industries 
Dover, Del. 

Radio Corp. of America 
Van Nuys, Calif. 

Sanders Associates 
Nashua, N.H. 

Brown Engineering 
Hunts vi lie, Ala. 

Reynolds, Smith and Hill 
Jacksonville, Fla. 

Ingalls Iron Works 
Birmingham, Ala. 

Smith/Ernst (Joint Venture) 
Tampa, Fla. 
Washington, D. C. 

Power Shovel, Inc. 
Marion, Ohio 

Hayes International 
Birmingham, Ala. 

Bendix Aerospace Systems 
Ann Arbor, Mich. 

Aerojet-Gen* Corp. 
El Monte.. Calif. 



Instrument Unit 



Guidance Components for Instru- 
ment Unit (Including ST-12*IM 
Stabilized Platform) 

Communications and Instru- 
mentation Support, KSC 

Launch Operations/Complex 
Support, KSC 

Facilities Engineering and 
Modifications, KSC 

Portable Life Support System; 
LM ECS 



>Mi4 



Space Suits 



110A Computer - Saturn Checkout 



Operational Display Systems 
Saturn 

Discrete Controls 



Engineering Design of Mobile 
Launchers 

Mobile Launchers (ML) 
(Structural Work) 

Electrical Mechanical Portion 

of MLs 



Transporter 

Mobile Launcher Service Arms 

Apollo Lunar Surface Experi- 
ments Package (ALSEP) 

Service Propulsion System Engine 

NASA-KSC APR/70